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 CONFIDENTIAL
WP-TECHNICAL NOTE 55 -298

t.(-" I (UNCLASSIFIED TITLE)

WAR EMERGENCY THRUST AUGMENTATION

FOR THE J47 ENGINE INTHE F- 86 AIRCRAFT

WILLIAM A. DAILEY, Ist Lt, USAF

POWER PLANT LABORATORY

AUGUST 1955

jUN 5Sll'
CENTER
WRIGHT AIR DEVELOPMENT

55WCLPR-10637 - 45

CONFIDENTIAL

AF.wp.-o-31 MA Y 6
NOTICE: THIS DOCUMENT CONTAINS INFORMATION AFFECTING THE

NATIONAL DEFENSE OF THE UNITED STATES WITHIN THE MEANING

OF THE ESPIONAGE LAWS, TITLE 18, U.S.C., SECTIONS 793 and 794.

'THE TRANSMISSION OR THE REVELATION OF ITS CONTENTS IN

ANY MANNER TO AN UNAUTHORIZED PERSON IS PROHIBITED BY LAW.

. -
 CONFIDENTIAL

WADC TECHNICAL NOTE 55- 298

(UNCLASSI FIED TITLE)

WAR EMERGENCY THRUST AUGMENTATION

FOR THE J47 ENGINE IN THE F.-86 AIRCRAFT

William A. Dailey, .1st Lt, USAF

Power Plaut Laboratory

A ugust 1955

Project No. ,T - 11 - P - 206A - 16

Wright Air Development Center

Air Research and Development Command
United States Air Force
Wright-Patterson Air Force Base, Ohio IJUN 1 5 1956

CONFIDENTIAL
 CONFIDENTIAL
FOREWORD

This document was prepared to serve as the final

report of Power Plant Laboratory Project T-11-P-206A-16,
(formerly S506-224J) entitled (Unclassified) "War Emer-
gency Thrust Augmentation for the J47 Turbojet Engine
Installed in the F-86 Aircraft". The project was adminis-
tered by the Rotating Engine Branch, Power Plant Laboratory,
Wright Air Development Center.
Acknowledgment for their work on the project is given
to the following: F-86 Weapon System Project Office,
Wright Air Development Center; Air Force Flight Test Center,
Edwards Air Force Base; NACA Lewis Flight Propulsion
Laboratory; General Electric Company; and North American
Aviation Inc.

This document, excepting the title and Sections I and

II, is classified CONFIDENTIAL because of the nature of,
and potential future military application of the work re-
ported on; classification is made in accordance with Air
Force Regulation 205-1, paragraph 24a, dated 15 December
1953.

WADC TN 55-298

CONFIDENTIAL 5 -10637
 CONFIDENTIAL
ABSTRACT

Augmentating the thrust of the J47 engine in the F-86

aircraft was the principal objective of the program herein
reported upon. The work covered a period of approximately
two and one-half years and was conducted primarily by the
General Electric Company and North American Aviation, Inc.
under Air Force contract. Work was also accomplished on
the project by the Wright Air Development Center, the Air
Force Flight Test Center, and the Lewis Flight Propulsion
Laboratory. All known schemes that could possibly augment
the thrust of a turbojet engine were considered and over-
speed, overtemperature, liquid nitrogen injection, water-
alcohol injection, and pre-turbine fuel injection were
brought under development. With the exception of overspeed
and liquid nitrogen injection, the above systems were
tested in flight and demonstrated that they could provide
increased thrust for the J47 engine thereby substantially
increasing the performance of the F-86 aircraft. Of the
three augmentation systems flight tested, two drastically
reduced the life of the engine and the third, water-alcohol
injection, although not having such a severe effect on en-
gine life was not suited for installation in the F-86 air-
craft. Thus, under the circumstances, adaptation of any
thrust augmentation system to the J47 engine in the F-86
aircraft for Korean operational use was deemed impractical.

PUBLICATION REVIEW

This report has been reviewed and is approved.

FOR THE COMMANDER:

Colonel, UTSAF
Chief, Power Plant Laboratory
Directorate of Laboratories

WADC TN 55-298 iii

55MCLPR-310637
CONFIDENTIAL
 CONFIDENTIAL
TABLE OF CONTENTS

Section I Overspeed 1

A. General-Testing-Results 1

Section II Overtemperature 3

A. General 3
B. Testing 4
C. Results 6

Section III Liquid Nitrogen Injection 13

A. General 13
B. Testing 14
C. Results 17

Section IV Water-Alcohol Injection 21

A. General 21
B. Testing 22
C. Results 27

Section V Pre-Turbine Injection 33

A. General 33
B. Testing 34
C. Results 41

Section VI General Conclusions 49

References 52

Appendix 54

WADC TN 55-298 iv

C I T
 CONFIDENTIAL
LIST OF ILLUSTRATIONS

Figure Page

1. Project's Significant Events and Dates x

2. J47-GE-13 Engine Thrust Vs RPM For Various

Tail-Pipe Area Settings 2

3. J47-GE-13 Engine Compressor Efficiency

Drop-Off With Increased RPM 2

4. F-86L Tail-Pipe Restrictor Segments Utilized

In the J47-GE-13 Engine Overtemperature Tests 5

5. Auxiliary Throttle Stop Installed On the Throttle

Quadrant In the Cockpit Of the F-86E Aircraft
During the J47-GE-13 Engine Overtemperature
Tests 5

6. Overtemperature Test Rate of Climb Data 7

7. Overtemperature Test Time to Climb Data 7

8. Overtemperature Test Level Flight True

Air Speed Data 8

9. Overtemperature Test Engine Net Thrust Data 8

10. Overtemperature Test Specific Air Range Data 10

11. Overtemperature Test Engine Gross Thrust Data 10

12. Schematic Diagram Of the Liquid Nitrogen Test

Set Up Utilizing the J47-GE-15 Engine 16

13. Liquid Nitrogen Injection Test Percent Augmen-

tation Vs Injection Flow Rate 18

14. Liquid Nitrogen Injection Test Inlet Air Flow

Temperature Decrease Vs Injection Flow Rate. 18

WADC TN 55-298 v

"CONFIDENTIAL
 CONFIDENTIAL

15. Schematic Diagram Of the J47-GE-27 Engine

Water-Alcohol Injection System Installed In
the F-86F Aircraft 23

16. J47-GE-27 Engine Water-Alcohol Injection

Control Panel On the Left Forward Console
In the Cockpit of the F-86F Aircraft 25

17. F-86F Tail-Pipe Tab Assembly Utilized In

the J47-GE-27 Engine Water-Alcohol Injection
Testing 25

18. Water-Alcohol Injection Test Rate Of Climb

Data 28

19. Water-Alcohol Injection Test Time To Climb

Data 28

20. Water-Alcohol Injection Test Level Flight

True Air Speed Data 30

21. Water-Alcohol Injection Test Engine Net

Thrust Data 30

22. Water-Alcohol Injection Test Aircraft

Acceleration Data 31

23. Water-Alcohol Injection Test Injection

Flow Rate Data 31

24. Schematic Diagram Of the J47-GE-27 Pre-

Turbine Fuel Injection System Installation
In the F-86F Aircraft 36

S25. F-86F Tail-Pipe Variable Area Nozzle

Utilized In the J47-GE-27 Engine'Pre-
Turbine Fuel Injection Tests 39

26. J47-GE-27 Engine Pre-Turbine Fuel. Injec-

tion Control Cockpit Presentation In the
F-86F Aircraft 39

27. Pre-Turbine Fuel Injection Test Rate Of

Climb Data 42

WADC TN 55-298 vi

CONFIDENTIAL
 CONFIDENTIAL

28. Pre-Turbine Fuel Injection Test Time To Climb

Data 42

29. Pre-Turbine Fuel Injection Test Level Flight

True Air Speed Data 43
30. Pre-Turbine Fuel Injection Test Engine Net
Thrust Data 43

31. Pre-Turbine Fuel Injection Test Aircraft

Acceleration Time Data 44

32. Pre-Turbine Fuel Injection Test Specific

Fuel Cornsumnpt•.on Dai,. 44,4

33. Pre-Turbine Fuel Injection Test Aircraft

Load Factor Ddta 46
34. Pre-Turbine Fuel Injection Test Radius Of
Turn Data 46
35. MiG-15 Rate Of Climb Data 54

36. MIG-15 Time To Climb Data 54

WADC TN 55-298 vii

CONFDENTIAL
 CONFIDENTIAL
INTRODUCTI ON

Encounters against the enemy MiG-15 aircraft in Korea

dictated that steps be taken to improve the performance of
the F-86 aircraft. Combat experience demonstrated the in-
feriority of the combat and service ceilings of the F-86 air-
craft and it was evident that the aircraft's climb perform-
An!e was in need of the most improvement. Although the F-86
aircraft had higher level flight and dive speeds than the
MiG-15 aircraft, the latters ability to accelerate more
rapidly to its maximum speed in part made up for its some-
what lower maximum attainable speed. It was directed that
the greatest effort be exerted on measures to improve the
combat capability of the F-86 aircraft with regard to the
above mentioned aircraft performance variables. The pro-
duct of these measures was to be something that could be
placed in operational combat use at the earliest possible
date; however, work was to be continued to provide further
Simprovements as soon as they could be made available to
supplant or supplement initial expediencies adopted as emer-
gency solutions. The understanding was that measures
which promised significant performance gains would not be
rejected on the grounds that engine life would be reduced
unless the reduction in life imposed an unsupportable bur-
den on supply and maintenance.
The deficiencies of the F-86 aircraft when compared to
the MiG-15 aircraft, as outlined above, were indicative of
one factor which needed improvement-thrust loading. In
order to make the thrust loading of the F-86 aircraft more
nearly equal to that MiG-15 aircraft, an improvement of such
in the order of 50% would have been necessary. The problem
was attacked by attempting to increase the thrust of the F-86
aircraft and reduce the aircraft's weight. All the known
methods of augmenting the thrust of a turbojet engine were
considered and many were actually brought under development
for the J47 engine, the purpose being to improve the perform-
ance of the F-86 aircraft. This report will present and re-
view the various technical aspects of the above cited problem.
which concern the J47 engine.
Because of the varied and extensive nature of the pro-
gram concerning the efforts to augment the thrust of the

WADC TN 55-298 viii

CONFIDENTIAL
 CONFIDENTIAL
J47 engine, the work reported upon herein is presented
chronologically rather than in any order pertaining to the
relative importance of the various phases of the work
accomplished or the benefits derived from such work. As
the object of the entire project was to meet an emergency,
where timing was necessarily important, pertinent data
with regard to this are presented in Figure 1 for early
reference. Other steps taken to improve the performance
of the F-86 aircraft, which were not directly related to
the J47 engine, such as rocket boost, aerodynamic improve-
ments, weight reductions, etc. are reviewed in numerous
other USAF and contractors' documents and will not be sub-
jects of this report.

WADC TN 55-298 ix

CONFIDENTIAL
 CONFIDENTIAL

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'WADC TN 55-296
CONFIDENTIAL
 CONFIDENTIAL
SECTION I

OVERSPEED

A. General - Testing - Results

Operating at a higher rotor speed was considered as a

means of increasing the thrust of the J47-GE-13 engine; it
was originally thought that an increase in thrust could be
obtained by increasing the air mass flow through the en-
gine. When the exhaust nozzle area was not enlarged for
such operation, an overtemperature condition existed. How-
ever, it was soon found that when the exhaust nozzle aroa
was so enlarged to maintain the limiting exhaust gas tem-
perature, a decrease in thrust, rather than an increase
occurred. Figure 2 illustrates how the thrust tapers off
if the exhaust gas temperature is maintained constant by
increasing the exhaust nozzle area for rotor speeds great-
er than 7950 rpm (100%). This condition was due to the
comparatively large decrease in the pressure ratio across
the nozzle relative to the increase in the air mass flow
through the engine. Overspeed operation in itself con-
tributed nothing to any additional thrust output when such
operation was combined with overtemperature due to fall-
off in compressor efficiency at the higher values of rpm;
Figure 3 illustrates this fall-off in compressor effi-
ciency. Should overspeed operation of the engine have
resulted in an increase in thrust, it is doubtful if such
operation could have been approved for any extensive ser-
vice use because of the relatively large increase in cen-
trifugal stresses which accompained overspeed. These
stresses increased approximately 8% when the engine was
oversped by 4%.

WAPO TN 55-298 1 551'ICLPR-10637

CONFIDENTIAL

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 CONFIDENTIAL
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WADO TN 55-298

CONFIDENTIAL
 CONFIDENTIAL
SECTION II

OVERTEMPERATURE

A. General
Previous experience had demonstrated that any thrust
increase which was derived during the engine overspeed
testing was solely due to overtemperature. Overtemperature
operation, without overspeed, was then considered as &
means of increasing the thrust of the J47-GE-13 engine.
Although it was known that overtemperature operation would
provide a substantial amount of augmentation, dependent
upon the degree of overtemperature, there were many prob-
lems which had to be resolved before such operation could
be approved for use in the field. The most difficult of
these problems was that of determining a life for the en-
gine parts that had been subjected to such overtemperature.
It was known from past experience that realistic engine
partst life figures could only be obtained through service
use or a long endurance testing program. Possible trouble
areas which needed investigation were as follows: ade-
quacy of aircraft cooling in parts adjacent to the engine,
-ability of the airframe structure to take the additional
loading and temperatures, and the determining of accurate
means of setting the rpm and exhaust nozzle area of the
engine on the ground to give the desired conditions at
altitude. Attention needed also to be given toward re-
solving the problems associated with the use of a small-
er area exhaust nozzle with regard to engine control opera-
tion, engine acceleration, engine stalls, flame-outs, and
starting.

Since it was not possible to enter into an endurance

type of test program to gather engine parts' life data, it
was decided to determine aircraft performance utilizing
overtemperature operation of the engine and to gather as
much other information as could be had, coincident with
such performance testing, that might allow a reasonable
assessment to be made about the degree of maintenance and
logistic support which would be necessary to allow tactical
use of engine overtemperature operation.

WADC TN 55-298 3

CONFIDENTIAL
 CONFIDENTIAL
B. Testing

Only a small expenditure of work was necessary to allow

overtemperature operation and that consisted of placing
restrictor segments in the tail-pipe to decrease the ex-
haust nozzle area. Figure 4 illustrates the restrictor
segments placed in the tail-pipe of the J47-GE-13 engine
in the F-86E test aircraft. In addition, an auxilial'y stop
was placed on the throttle quadrant in the cockpit. The
auxiliary stop was installed on the throttle quadrant so
that the engine might not be continually operated at an
overtemperature condition. The auxiliary stop was set at
a position which limited the rpm of the engine to a value
such that 100% thrust was obtained at 100% exhaust gas
temperature but at lower engine rpm than 100%. The auxil-
iary stop was designed so that the throttle could be pushed
outboard and forward to the original full throttle position
thus giving additional thrust resulting from overtempera-
ture operation. Figure 5 illustrates the auxiliary stop
installed on the throttle quadrant in the cockpit of the
F-86E test aircraft. Such a system allowed normal opera-
tion up to the auxiliary stop and in addition allowed the
pilot to obtain additional thrust by advancing the throttle
past the auxiliary stop. Since the rpm was also increased
when the throttle was advanced past the auxiliary stop,
the air mass flow through the engine was increased slightly.

Approximately 99% thrust was available with the throt-

tle at the auxiliary-stop provided 100% exhaust gas tem-
perature was maintained at that setting. Not only was 99%
thrust available at 93% rpm, but due also to the engine's
component characteristics, the full thrust normally avail-
able was obtained between 96% and 100% rpm at 100% exhaust
gas temperature with due allowances being made for changes
away from standard ambient conditions. Static and flight
testing was accomplished with the auxiliary stop set at
7400 rpm (93%) and the exhaust nozzle area so reduced to
produce 12750F(100%) exhaust gas temperature at that en-
gine rpm. With such an arrangement, an exhaust gas tem-
perature of 1500OF (118%) was produced at 7950 rpm (100%).
A timer was installed so that the duration of operation
above 93% rpm could be determined. Flight operation above
93% rpm was for the most part limited to one minute cycles
at the overtemperature setting in order to obtain maximum

WADC TN 55-298 4
CONFIDENTIAL
 CONFIDENTIAL

FIGURE 4. F-86E Tail-Pipe Restrictor Segments

Utilized In The J47-GE-13 Engine
Overtemperature Tests.

PIGURE 5. Auxiliary Throttle Stop Installed, On The

Throttle Quadrant In The Cockpit Of The
F-86E Aircraft During The J47-GE-13 Engine
Overtemperature Tests

WADC TN 55-298 5
.9 CONFIDENTIAL
V

CONFIDENTIAL

life of the turbine buckets. Continuous and cyclic ovArtem-

perature operation tests were conducted both on the ground
and in flight. Throughout the testing, inspections were
made in an attempt to correlate parts' life deterioration
with overtemperature operation. A.full metallurgical in-
vestigation of all parts subjected to overtemperature was
made upon completion of the testing.
Two other tests were run in conjunction with the over-
temperature flight testing. One such test was the placing
of a by-pass needle valve between the large and small slot
fuel manifolds to prevent exhaust gas temperature drop-off
in climbs and the other was the placing of a 30 psi restric-
tor valve in the emergency fuel system side of the double
check valve so as to give better temperature regulation
while climbing. Both of these additional tests were con-
ducted with the purpose of alleviating problems which had
been experienced in the field.

-C. Results
Figures 6 through 11 are plots of pertinent perform-
ance data gathered during the testing phase of an P-86E
aircraft with a J47-GE-13 engine utilizing overtemperature
as a means of thrust augmentation. Data for a standard
unaugmented J47-GE-13 engine and F-86E aircraft are sup-
plied for comparative purposes.
From Figure 6 it can be seen that the rate-of-climb
of the F-86E aircraft wps increased by 1800 ft/min at alti-
tudes between 30,000 feet and 45,000 feet when utilizing
overtemperature operation of the engine; thus the rate-of-
climb was nearly doubled at an altitude of 35,000 feet and
nearly tripled at an altitude of 40,000 feet. At an alti-
tude of 45,000 feet the rate-of-climb with overtemperature
operation was four times as great as the rate-of-climb at
military power. The time to climb from an altitude of
30,000 feet to an altitude of 45,000 feet wasnaeduced by
approximately 8 minutes utilizing overtemperature opera-
tion of the engine as can be seen from Figure 7. Figure 8
shows a steady incremental increase in maximum level flight
true air speed of about 10 knots at an altitude of 15,000
feet to 15 knots at an altitude of 45,000 feet utilizing
overtemperature operation as compared to operation at mili-

WADC TN 55-298 6

CONFIDENTIAL
* r
 CONFIDENTIAL

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WADO TN 55-298 7

CQNýfIDETIA
 CONFIDENTIAL

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WADC TN 55-298 8

CONID-ENTIAL
 CONFIDENTIAL

tary power. The performance gains as noted were made possi-

ble by an average increase in net thrust of approximately
27% at altitudes between 15,000 and 45,000 feet, with a
maximum increase in net thrust of nearly 30% being obtained
at an altitude of 35,000 feet as can be observed from Fig-
ure 9.

As a result of the smallet exhaust nozzle area, a

lower specific fuel consumption was obtained at the inter-
mediate power settings within the cruise regime which was
below 93% rpm. At the 100% rpm position, utilizing over-
temperature operation, a somewhat lower or at least equal
specific fuel consumption was realized. As detailed pre-
viously, no overtemperature operation was allowed below
93% rpm; 93% rpm was the point at which the auxiliary
throttle stop was placed. Figure 10 shows a favorable
increase in the range obtainable for the F-86E aircraft
utilizing the smaller exhaust nozzle area necessary for
overtemperature operation; the data indicates a 4 to 8%
increase in range for a given fuel load, the cruising
altitude naturally varying with aircraft weight. The
curves presented in Figure 10 were obtained for best cruise
speeds based upon calculation. Estimates indicated that
an increase in combat radius of approximately 20 nautical
miles was possible for the basic mission if flown with the
non-standard exhaust nozzle area setting utilized for
overtemperature operation.
As was stated previously, the deterioration of engine
life was an important factor to be evaluated in the program.
An accumulation of parts' life data allowed the following
accessment of life to be given to each part: turbine buck-
ets -- 10 minutes, transition liners, turbine nozzle dia-
phragm, and shroud ring -- 20 minutes, and inner combustion
chambers -- 30 minutes. The above parts' life figures re-
fer to overtemperature operation in a cumulative sense and
represent the life of the parts for an increase in tempera-
ture of 225oF from the standard value. The exhaust-cone
and tail-pipe were also adversely affected by the overtem-
perature operation- exhaust cone cracking was quite common.
The results of the static tests showed that cyclic opera-
tion at overtemperature and continuous operation to the
same total overtemperature level and duration were equally
bad from a parts' life viewpoint.

WADC TN 55-298 9
CONFIDENTIAL
 CONFIDENTIAL

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,"WADC TN 55-298 10

CONFIDENTIAL
 CONFIDENTIAL

The ten minute life of the turbine buckets was found

to be a-limiting factor. In service use it would be nec-
essary for all the buckets to be replaced at that time for
it would not be possible to determine visually If the buck-
ets were fit for further service. It was felt that the
remaining parts of the engine subjected to overtemperature
could be given a standard hot parts inspection and allowed
to remain in service provided all the inspection criteria
established in the already existing technical orders for
such inspections were met. The.effect of overtemperature
on the turbine wheel was only determined to avery limited
extent due to the difficulty surrounding such an investiga-
tion. In order to utilize overtemperature operation for
the maximum time limit as determined by the life of the
turbine buckets, it was necessary to begin the testing with
new buckets. It was verified from the tests that if the
overtemperature system as evaluated was to be used in the
field with a reasonable factor of safety, new turbine
buckets would have to be installed at the beginning of
overtemperature operation and replaced after 10 minutes of
overtemperature operation had been accomplished. Coinci-
dent with this turbine bucket replacement a thorough hot
parts inspection was felt necessary. Such a bucket re-
placement and inspection task would necessitate removal
of the aft fuselage of the aircraft and removal of the
turbine wheelfrom the engine.

In general, the entire system functioned properly and

no outstanding engine control or aircraft cooling problems
were encountered. The test results indicated that the in-
crease in performance of the F-86E aircraft resulting from
overtemperature operation of the J47-GE-13 engine was sub-
stantial but the maintendnce and logistic support for such
operation made use of engine overtemperature operation im-
practical; thus, no overtemperature as such was utilized in
Korea. A variation of the previous described technique
used during the overtemperature tests was employed by one
-fighter wing in Korea, the purpose being to prevent exhaust
gas temperature drop-off with altitude. The exhaust nozzle
was tabbed by the placing of restrictor segments in the
tail-pipe to produce rated temperature (100%) at 96% rpm
while on the ground and as the temperature dropped off with
altitude, the throttle was advanced by the pilot. Such a
system if properly used did not overtemperature the engine,
but it was necessary that the pilot closely monitor the
WADC TN 55-298 11
CONFIDENTIAL
 ;I'

CONFIDENTIAL
exhaust gas temperature. The by-pass needle valve between
the large and small slot fuel manifolds, which was evalu-
ated during a portion of the overtemperature testing, was
also used in Korea by another fighter wing to accomplish
the same task of reducing exhaust gas temperature drop-off
with altitude. Reduction in the temperature drop-off made
possible the realization of more nearly the full available
thrust at altitude and provided the limiting temperature
was not exceeded, there would be no reduction in parts'
life. As was previously stated, no loss in engine thrust
occurred at the slightly reduced rpm, which was used for
take-off and operation at the lower altitudes,provided
100% exhaust gas temperature was maintained.

WADO TN 55-298 32

CONFIDENTIAL
"-I
 CONFIDENTIAL
SECTION III

LIQUID NITROGEN INJECTION

A. General

One method of thrust augmentation for the J47 engine

that appeared promising on the basis of analytical work
conducted was that of liquid refrigerant injection into
the compressor inlet. Such injection artifically cooled
the entering air thereby resulting in a greater air mass
flow through the engine and hence increased thrust. The
lower temperature of the air flowing through the compres-
sor increased the compressor pressure ratio by increasing
the compressor Mach Number and also increased the differ-
ence between the temperature at which the work of compres-
sion was added to and taken from the working fluid in the
engine. Liquid nitrogen and liquid oxygen were the refrig-
erants given the most consideration in the studies. It was
concluded from a theoretical examination that for the same
weight flows of refrigerant injection, liquid nitrogen and
liquid oxygen would give approximately the same thrust aug-
mentation providing the combustion process was not serious-
ly disturbed by the injection of either refrigerant. Li-
quid oxygen is however more dense than liquid nitrogen and
p has the advantage of requiring a 14% smaller, storage tank
for the same weight. The use of liquid oxygen was later
ruled out because it is dangerous from a handling point of
view; it may explode spontaneously when brought in contact
with grease or oil. Experience had also shown that leak-
ages in installations utilizing liquid oxygen do occur and
that the fire hazard is considerable; nitrogen on the other
hand is inert and does not burn or explode.
Hydrogen peroxide, liquid ammonia, methyl chloride, and
liquid air were also considered for possible use but soon
eliminated from further consideration. Hydrogen peroxide
was disqualified as a satisfactory coolant at the compara-
tively low operating temperatures since its boiling tem-
perature is relatively high. Hydrogen peroxide also pre-
sented a distinct problem for it is spontaneously combus-
tible at temperatures approximating the compressor dis-
charge temperature of the J47 engine. Liquid ammonia,

WADC TN 55-298 13
CONFIDENTIAL

ii
 CONFIDENTIAL

although possessing a boiling temperature sufficiently low

to permit favorable evaporation at the temperatures and
pressure considered is' nevertheless toxic and since com-
pressor discharge air was used for aircraft cabin pressuri-
zation, its use was ruled out. Liquid ammonia also attacks
copper and copper alloys in the presence of moisture. Me-
thyl chloride and liquid air were considered unsuitable be-
cause their heats of vaporization are comparatively low,
and thus these coolants could absorb only small quantities
of heat during vaporization. Water injection into the com-
pressor inlet was unsatisfactory for use with the J47 en-
gine due to the cooling of the compressor case and its
subsequent contraction which caused interference between
the rotor blades and the case. Naturally, cooling of the
engine's compressor case would also occur with any type of
refrigerant used, but it was thought that by injecting the
coolant near the very beginning of the duct leading to the
engine instead of directly at the compressor face, the
throwing of the coolant outward toward the case as a result
*of centrifugal action could be avoided. It was doubtful
if water injection into the compressor inlet would have
been of much benefit at altitude because although water is
a satisfactory coolant at the normal air temperatures asso-
ciated with near sea level operation or for very high
speeds at the higher altitudes, it is unsatisfactory at
the low temperatures encountered by aircraft operating at
moderate speeds at altitude since small amounts of water
saturate the air and very little evaporate cooling can be
obtained.

Since liquid oxygen appeared to offer no advantages

as compared to liquid nitrogen injection, but instead offered
many disadvantages in the elaborate care and precautions
required for its safe handling and use, a test program was
begun using liquid nitrogen. The purpose of initiating
such a program was to determine whether satisfactory opera-
tion of the J47 engine could be maintained with liquid
nitrogen when it was injected into the compressor inlet and
also to determine the amount of thrust augmentation pro-
duced by such injection.

B. Testing

Only static engine testing was accomplished with four-

teen actual test runs being made with liquid nitrogen in-
WADC TN 55-298 14
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t
 CONFIDENTIAL

jection. Figure 12 is a diagram of the test set up. Two

tanks were used, one was a large supply tank and the other
was a smaller pressure tank. To minimize heat transfer
and the resulting loss of liquid nitrogen, the two tanks
and the related plumbing were wrapped with an insulating
foil. Gaseous nitrogen was used to pressurize the liquid
nitrogen supply tank to force the liquid through the in-
jection nozzle. A manually controlled pressure regulator
was used to control the flow of gaseous nitrogen into the
liquid nitrogen tank and the resulting liquid through the
injection nozzle. An F-86 aircraft duct was placed ahead
of the engine and was separated from it by a plenum chamber.
The injection nozzle was located at the very beginning of
the duct and extending directly into the center, being sup-
ported by its single supply line which extended out from
one side of the duct. The position of the liquid nitrogen
injection nozzle was approximately 24 feet linear horizon-
tal distance from the plane of the engine compressor inlet.
The nozzle stem axis was parallel to the axis of the engine.
The geometry of the nozzle was'such that the nitrogen was
injected in a continuous sheet which in still air would
form a cone with an included angle of approximately 100 de-
grees. The geometric apex of the cone pointed downstream
so that there was an upstream component of injection velo-
city. However, when the engine was running at near full
power, the velocity of the entering air overcame the up-
stream component of the nitrogen injection velocity so that
the injected sheet of liquid nitrogen took on the appear-
ance of a paraboloid of revolution with its vertex located
at the injection nozzle with the concave formation opening
downstream. Originally it was felt that the smallest
practical nozzle orfice size and the largest practical
nozzle injection pressure would provide the optimum vapori-
zation because such a combination would provide the best
atomization. Optimum vaporization was desired because if
the liquid nitrogen was not completely vaporized before
reaching the compressor inlet, a portion of the potential
cooling effect on the engine airflow, and consequently
maximum thrust augmentation would not be realized. It was
later found that another factor influenced the amount of
vaporization. It appeared that with a greater nozzle open-
ing, larger slugs of liquid nitrogen were injected into
the engine airflow and it was reasoned that the increased
mass and added momentum were such that it resulted in the

I:L
WADC TN 55-298 1i
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 CONFIDENTIAL

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WADC TN 55-298 16

CONFIDENTIAL
 CONFIDENTIAL
liquid nitrogen penetrating the engine airflow all the way
to the duot walls where the nitrogen was spread out in a
thin film and vaporized by the agitating nature of the
boundary layer. Thus better results were obtained by using
a larger nozzle opening than was originally thought neces-'
sary.
To offset the exhaust gas temperature drop-off during
liquid nitrogen injection, the same method previously de-
scribed was utilized; thus prior to the test, the exhaust
nozzle area was so adjusted to produce the maximum allow-
able continuous exhaust gas temperature at approximately
93% rpm. The throttle was then advanced during injection,
usually to approximately 97% rpm in order to maintain the
exhaust gas temperature near its maximum allowable limit.

C. Results
Figure 13 shows the percent thrust augmentation ob-
tained for various liquid nitrogen injection rates utiliz-
ing the J47 engine in static tests. It can be seen that
at the relatively high liquid nitrogen injection rates in
the order of 17 lbs/sec an increase in thrust augmentation
of approximately 28% was obtained. There was quite a dis-
crepancy between the data obtained from preliminary calcu-
lation and the actual test results; it was believed to be
caused by the poor vaporizing ability of the injection
nozzle. Thus a smaller increase in actual inlet weight
flow and compressor pressure ratio were obtained which re-
asultdd in a lower thrust due to the non-uniform inlet-air
temperature distribution. It appeared that the larger the
nozzle openings and the higher the injection rates, the
worst the engine inlet temperature distributions were. As
was previously pointed out, the utilization of a somewhat
larger nozzle than was originally anticipated had also a
rather beneficial effect.
Figure 14 shows the approximate maximum temperature
decrease of the airflow entering the engine as a result of
liquid nitrogen injection. The data presented in the figure
were obtained after the injection flow had built up to a
relatively constant value. The approximate point to point
temperature reduction of the airflow across the compressdr
inlet showed a i 10% to ± 14% variation from the mean value
with liquid nitrogen injection. The total pressure loss
of the incoming air to the engine due to the injection of

WADC TN 55-298 17

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 CONFIDENTIAL
liquid nitrogen was about twice as much as was indicated by
preliminary calculation. It was felt that the higher than
anticipated pressure loss could be attributed to the fact
that the nitrogen was injected at an angle having an up-
stream velocity component. A loss in free stream total
pressure of 27due to nitrogen injection would, by calcula-
tion, amount to a net thrust decrement of about 115 lbs for
the J47-GE-13 engine in the F-86 aircraft at an altitude
of 30,000 feet and a true air speed of 550 knots. Thus
pressure loss was a more important factor than theoritical
analysis had indicated. However,it is felt that additional
nozzle development could minimize the total pressure loss
due to injection. Engine combustion failure or flame-out
occurred in 40% of the tests under similar conditions. It
is believed that they were precipitated by excessive rates
of liquid nitrogen injection in excess of 17 lbs/sec. It
is significant then that the flame-outs occurred when the
compressor inlet total temperature was in the region of
4000R.
In handling the liquid nitrogen during the tests, the
fact became generally established that liquid nitrogen was
not nearly as volitile as some references had pointed out.
The nitrogen tanks were prone to leaking at any place there
was a bolt in a hole due to the cooling effect on the met-
als and the differential contraction between the two. Al-
though the handling of the liquid nitrogen appeared quite
reasonable, its storage could prove quite difficult and
its availability might be limited as a result; another fac-
tor was the excessive weight and space necessary for insu-
lating aircraft storage tanks. No data was collected on
the effect that liquid nitrogen injection into the engine
would have upon the cabin pressurization equipment since no
attempt was made to adapt the system to an F-86 aircraft.
It was thought that through continual development of the
system, a substantial increase in thrust could be gained
even at the higher altitudes; but, it was also thought that
the limitations imposed on the system from an aircraft
modification and weight standpoint made it impractical.
There was also the time that had to be made available to
acquire flight test data on such a system and it was be-
lieved that concentration should be centered on a system
that showed promise of being more easily adapted.
Liquid nitrogen injection was eliminated from further
consideration as a means of augmenting the thrust of the
WADC TN 55-298 19
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 CONFIDENTIAL

J47 engine in the P-86 aircraft even though static testing

and estimated performance at altitude had showed promise.
It was nevertheless true that the full potential of such a
system could only be realized at relatively high ambient
temperatures or very high speed operation and such which
would not be the case for operation in Korea with the F-86
aircraft against the enemy MIG-15 aircraft.

WADC TN 55-298 20

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SECTION IV

WATER-ALCOHOL INJECTION

A. General
Although analytical work showed that liquid injection
into the compressor might prove highly satisfactory, such
operation with the J47 engine was not possible due to the
cooling and subsequent contraction of the compressor case
causing interference between the case and the rotor blades;
therefore, the only alternative was to inject directly in-
to the combustion chambers. Work on a wate:'-alcohol com-
bustion chamber injection system for the J47-GE-13 engine
in the F-86 aircraft was therefore initiated. The basic
idea behind the water-alcohol combustion chamber injection
system as applied to the J47 engine was that by virtue of
the liquid injection, the fluid weight flow through the
engine could be increased. In addition, the exhaust gas-
pressure would also increase. Both of these factors al-
lowed increased thrust. It was necessary to mix alcohol
with the water so as to supply the heat required for va-
porization of the water. Some work on such a system had
already been accomplished with the J47 engine prior to the
initiation of a formal program to meet the then present
emergency, but it was confined to static testing.
There were many problems associated with the use of a
water-alcohol combustion chamber injection system in the
F-86 aircraft that had to be Investigated. Also, it was
not known prior to the initiation of flight testing just
exactly what increase in aircraft performance might be re-
alized with such a system. The problem of making a mechani-
cally satisfactory water-alcohol injection installation in
the F-86 aircraft was a difficult one since space was ex-
tremely limited. The actual components to be used in the
aircraft portion of the system presented problems for none
were specifically designed for such an installation. In the
interest of safety, the preliminary tests using water-alco-
hol injection were made on one engine of a B-45 aircraft.
Later testing was accomplished with a J47-GE-13 engine in
a F-86A aircraft. Final testing was accomplished with a
J47-GE-27 engine ina F-86F aircraft for the F-86E aircraft
was scheduled to be phased out of combat.

WADC TN 55-298 21

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B. Testing

The water-alcohoi mixtures utilized in the testing

consisted of plain tap water, AN-A-18 alcohol (MIL-A-6091),
and emulsive corrosion preventive oil, USAF Specification
3604-A. The mixtures were prepared on a volumetric basis.
At first, the mixtures were prepared by mixing the oil and
water, then adding this mixture to the alcohol. Later,
better results were obtained by adding the water to the
alcohol and mixing to the desired percentage, then adding
the oil. Figure 15 is a diagram of the water-alcohol
injection system tested in the F-86 aircraft. The air-
craft's aft fuselage 105 gallon fuel tank was isolated for
use as a water-alcohol tank. On the original installation
the normal tank outlet line was used as a water-alcohol
supply line. and the fuel transfer pump was replaced by a
pump which had been modified for use as a water-alcohol
boost pump. It was soon learned that the tank outlet line
offered too much restriction and the boost pump could not
*supply the flow required to keep the injection system in
operation at the lower altitudes, so the tank outlet line
was capped and use of the boost pump was abandoned.

For the next configuration a plate was made to fit in

place of the fuel level transmitter on top of the tank and
a 1-3/4 inch diameter tube was welded to the plate and
formed in such a way that it extended downward to within an
inch of the bottom of the tank. Air for tank pressuriza-
tion was obtained from the line used for pressurizing the
main hydraulic reservoir. A one-half inch diameter line
was installed between the tee downstream from the air-
craft's primary heat exchanger and one of the tank vent
lines. A gate type shut-off valve was installed in the line,
and a relief valve capable of passing high airflows was
used to limit maximum tank pressure to 8 psi. The-other
tank vent lines were capped. This configuration proved
very successful and was used for the remainder of the test
on the F-86A aircraft.

For the test on the F-86A aircraft, the J47-GE-13 en-

gine was equipped with thimble type combustion chamber
liners since it had been determined that the combustion
characteristics, when utilizing these liners, were better
when used in place of the standard liners. For approxi-

WADC TN 55-298 20

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mately half the test the standard 100 gal/hr water-alcohol

nozzles were used, and for the remainder of the test the
experimental 50 gal/hr nozzles were used with the lower
capacity pump.
The system was modified only slightly for the F-86F
aircraft test installation. In that aircraft, the ammuni-
tion compartment heating system was used to supply tank
pressurization. The connection to the system was made
just downstream from the heating air shut-off valve and
wiring was installed so the valve could be controlled from
the cockpit. A larger capacity turbine pump was installed
to overcome the difficulties encountered in the F-86A air-
craft installation. A plague of pump failures began,
apparently caused by cavitation which resulted in overspeed-
ing. Various modifications were made in a vain effort to
eliminate air from the water-alcohol pump inlet and outlet
lines. Eventually the water-alcohol supply line between
the tank and the pump was changed completely. A plate was
made to fit in place of the tank inspection door on the
aft face of the tank. A tube, welded to the plate, extend-
ed in to the center of the tank, and a mating line was
connected to the pump inlet. By keeping the line as low as
possible, trapped air was held to a minimum. In addition,
a valve was installed so that all air could be bled from
the line after each servicing; a slightly smaller capacity
pump was also installed to replace the large capacity pump.
The J47-GE-27 engine in the F-86F aircraft was not
normally equipped for water-alcohol injection. The J47-GE-
27 engine used for the testing was therefore modified by
installing an external water-alcohol manifold, eight flex
lines for connecting the manifold to each combustion cham-
ber and a set of combustion chambers from a J47-GE-25
engine which was used in a B-47 type aircraft. The water-
alcohol manifold encircled the forward end of the combus-
tion system and was attached to the compressor rear frame.
The experimental 75 gal/hr water-alcohol nozzles were used
throughout the test.
The pilot's water-alcohol injection control panel
consisted of two switches for opening and closing the tank
pressurization shut-off valve and the turbine-pump air con-
trol valve, appropriate circuit breakers, and a light which
indicated when the pressure switch closed; Figure 16 il-.
WADC TN 55-298 24
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 CONFIDENTIAL

0 S

FIGURE 16. J47-GE-27 Engine Water-Alcohol Injection

Control Panel On The Left Forward Console
In The Cockpit Of The F-86F Aircraft.

FIGURE 17. F-86F Tail-Pipe Tab i-issembly Tjtilized

In The J47-GE-27 Engine Water Alcohol
Injection Testing.

WADC TN 55-298 25
f't*M flD•tU
 CONFIDENTIAL
lustrates the water-alcohol injection control presentation
located in the cockpit. To initiate water-alcohol injec-
tion, the tank was first pressurized. After at least 15
seconds the water-alcohol injection control switch was
moved to PRIME and held, energizing the circuit to the air
control valve providing the float switch was closed. The
float switch was a safety feature installed to prevent
starting the pump until there was a head of water-alcohol
at the pump inlet. A air bleed valve was used in conjunc-
* tion with the float valve to allow air to escape. Pros-
surization of the water tank forced water into the float
valve assembly, raising the float and closing the switch.
and at the same time closing the bleed valve. When the
motor on the air control valve was energized, the valve
opened allowing compressor discharge air to energize the
turbine pump. As soon as pump discharge pressure was 10
psi greater than combustion chamber pressure, water-alco-
hol was forced through the check valve and injection was
started. Simultaneously the exhaust nozzle tabi which was
necessary to maintain temperature, was forced up into the
exhaust stream by the force of the pump discharge pressure
on the piston in the nozzle actuator. Figure 17 illustrates
the tab assembly used for the testing. As soon as the
water-alcohol pressure reached the pre-selected pressure
switch setting it closed the switch energizing the pilot's
indicator light and completed an alternate circuit to the
air valve.The pilot then released the switch and water-
alcohol injection coptinued until thepump discharge pres-
sure dropped to the level at which the pressure switch was
set to open. Opening of the pressure switch de-energized
the air valve and stopped the airflow to the pump. If the
pilot desired, he could stop water-alcohol injection by
moving his control switch to the OVERRIDE OFF position;
otherwise, the injection continued until the water-alcohol
mixture was expended.

In 35 flight hours on the test F-86A aircraft, approxi-

mately three (3) hours of water-alcohol injection time was
accumulated. For the test F-86F aircraft, in nearly 50
hours of flight testing, over three (3) hours of water-alco-
hol injection was accomplished. Many more hours of water-
alcohol injection were accumulated during ground tests and
some initial flight testing utilizing a B-45 aircraft was
accomplished.

WADC TN 55-298 26
CONFIDENTIAL
 CONFIDENTIAL
C. Results

In general, the first water-alcohol flight tests uti-

lizing a single J47 engine of a B-45 aircraft were success-
ful and most promising. A difficult task came however with
the adapting of the system to the F-86 aircraft.

The augmentation obtained with the configuration tested

in the F-86A aircraft utilizing a J47-GE-13 engine was
very encouraging, being approximately 30% at altitudes of
30,000 feet and above. No serious problems were encoun-
tered, and operation was for the most part quite satisfac-
tory at altitudes up to 40,000 feet. At higher altitudes
the termination of water-alcohol injection during climbs
invariably resulted in flame-outs, although level flight
operation was normal. Upon disassembly of a test J47-GE-13
engine for inspection after approximately 1-1/2 hours of
water-alcohol injection time, the combustion chamber inner
liners and transition liners were found to be damaged. The
damage was attributed to the poor and erratic spray pattern
exhibited by the standard water-alcohol nozzles in the
J47-GE-13 engine. During an equivalent period of operation
with lower flow capacity nozzles, no engine damage was in-
curred; however, the restriction of these nozzles was so
great that flow, and consequently augmentation, was notice-
.ably reduced even with maximum power Input to the pump.
Since a method of augmentation for the F-86F aircraft was
of primary interest, and because more intensive testing
was accomplished with that aircraft, further discussion
will be confined to that phase of the testing.

Figures 1? and 19 show comparative climb performance

data between a standard F-86F aircraft and one utilizing
water-alcohol injection. It can be observed from Figure
18 that the rate-of-climb of a water-alcohol augmented
F-86F is continually increased above an altitude of 25,000
feet until it is over double the dry rate-of-climb at
40,000 feet. As a consequence, the time to climb from an
altitude of 20,000 feet to an altitude of 30,000 feet
is reduced by approximately one minute. Also, the time to
climb from an altitude of 30,000 feet to an altitude of
40,000 feet is reduced by over four minutes. The reason
for the discontinuities in Figure 19 is that a change in
water-alcohol flow rate had to be made for each range of

WADC TN 55-298 27

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 CONFIDENTIAL
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WAflC TN 55-298 2

CONFIDENTIL.
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 CONFIDENTIAL
altitudes; such a change could only be made on the ground.
Figure 20 shows the increase in level flight true air
speed that was obtained by the use of water-alcohol injec-
tion;: it can be seen that the speed of the F-86F was in-
creased by 10 knots at an altitude of 20,000 feet and nearly
15 knots at an altitude of 45,000 feet. Maximum thrust
augmentation obtained in the F-86F aircraft varied from
approximately 19% at 20,000 feet to approximately 29% at
an altitude of 40,000 feet; such data are presented in
Figure 21. Since changing the water-alcohol injection rate
in order to extend operation to higher altitudes was a
ground adjustment, the reduction in flow necessary at high
altitude resulted in less flow and less augmentation at the
lower altitudes if the system was set for high altitude
operation. The augmentation at the lower altitudes was re-
duced in theorderof 20% so as to allow satisfactory opera-
tion at the higher altitudes and eliminate the need for an
adjustment. It was reasoned that satisfactory operation at
altitudes above 40,000 feet was worth the loss in augmenta-
tion which resulted at the lower altitudes. Figure 22
,shows comparative data gathered from accelerations from
minimum level flight true air speed to maximum level flight
true air speed at an altitude of 35,000 feet. It can be
noted that with water-alcohol injection, the F-86F aircraft
reaches the same maximum true air speed at an altitude of
35,000 feet that was possible with an unaugmented F-86F
aircraft approximately 1-1/4 minutes sooner.

More serious engine instability waa encountered with

*the J47-GE-27 engine than with the J47-GE-13 engine, and
alcohol ,percentages were more critical. Mixtures ranging
from 20O to 28% alcohol were used, but with the higher per-
centages there was a tendency for the engine to overspeed
when starting or stopping water-alcohol injection. The
maximum water-alcohol flow schedule used, as shown in Fig-
ure 23, which varied from 48 gal/min at an altitude of
20,000 feet to 33 gal/min at an altitude of 40,000 feet
gave optimum performance throughout that altitude range
with a 24% alcohol mixture. Above an altitude of 40,000
feet, that combination resulted in flame-outs when the
injection was terminated under conditions other than level
stabilized flight. By reducing the flow rate to 26 gal/mLin
at an altitude of 40,000 feet, satisfactory operation was
obtained during level flight, climbs, dives, and other

WADC TN 55-298 29

CONFIDENTIAL

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WADC TN 55-298 3
CONFIDENTIAL.
 CONFIDENTIAL

maneuvers at altitudes up to 48,000 feet; only one climb

was continued to an altitude of 50,000 feet. Operation
above 45,000 feet necessitated that the pilot more closely
monitor the exhaust gas temperature so as to keep it within
limits. Also at altitudes of 45,000 feet and above there
was usually surging in engine speed when water-alcohol in-
jection was initiated but throttle adjustment could allevi-
ate the problem. The single exhaust nozzle tab, as-located,
* caused an objectionable yaw force, but it was thought that
by a redesign or the symmetrically locating of twin tabs
would eliminate the trouble.
It was concluded that the normal combustion system
components of the J47-GE-27 engine have a water-alcohol in-
jection endurance life of approximately two hours. It is
noted from the forgoing discussion that water-alcohol aug-
mentation of the J47-GE-27 engine in the F-86F aircraft
offered substantial gains in performance. With the excep-.
tion of the 105 gallon fuel tank converted to carry water-
alcohol, the configuration tested was satisfactory; a
single aircraft under-fuselage tank was designedfalthough
never tested,to overcome that difficulty. It was apparent
however that some additional testing was necessary. It was
also apparent that even with ultimate refinement, the gross
weight of the aircraft would be considerably increased. One
problem tha4 appeared difficult in light of the fact that
operation was to be in Korea, was that of logistics, It
was decided that a met~wd'augmentation that did not require
another fluid and therefore not necessitate a duel tank
system would be the most desired, thereby, allowing flexi-
bility from mission to mission or as the need oceured during
any one mission. Since it was not possible to support a
high-priority effort on more than one project, it was
concluded that pre-turbine injection, which looked promis-
ing on the basis of preliminary analysis, would be concen-
trated on and that further work on water-alcohol injection
would be continued on a.development basis with direct appli-
eating being aimed at the B-47 aircraft for take-off only,
where the control problems would necessarily be less. It
was felt that a water-fuel injection system might be used.
With such a system JT-4 would be substituted for the alco-
hol.

WADC TN 55-298 32
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 CONFIDENTIAL
SECTION V

PRE-TURBINE INJECTION

A. General
One of the most popular means of augmenting the thrust
of a turbojet engine is afterburning. An afterburner for
the J47 engine had already undergone some development test-
ing as early as February 1948; this afterburner development
engine was designated the XJ47-GE-5. Further development
of that engine led to the J47-GE-17 engine which powers the
F-86D Aircraft and eventually to the J47-GE-33 engine. The
J47-GE-17 and the J47-GE-27 engine were by no means inter-
changeable. Even provided it would have been possible to
install a conventional afterburner in the F-86F aircraft,
the added pressure losses when non-afterburning would have
reduced the aircraft's cruising range. Since the cruise-
out portion of the missions in Korea were lengthy, the use
of a conventional afterburner would show a disadvantage
from that viewpoint.

A method of afterburning was necessary which would, in

addition to providing the increase in thrust necessary, also
be capable of being incorporated in the existing aircraft
with little modification and also be such that the dry en-
gine performance of the aircraft would not be affected.
With such rigid requirements, only pre-turbine injection
seemed feasible. Pro-turbine injection (hereafter referred
to as PTI) is a system of reheat whereby the fuel for after-
burning is injected upstream of the turbine in contrast to
the conventional method of injecting the fuel for after-
burning downstream of the turbine. With such a system the
turbine wheel acts as a flameholder rather than having
separate flameholders which contribute to the dry loss of
an afterburning engine. PTI, which in a liberal sense
might be referred to as an extremely short afterburner,
allowed the combining of the diffuser and burner sections
into one and therefore made installation in the already
existing F-86 aircraft possible. The first work on such
a system was conducted in Germany as early as October 1939.
The first successful tests were run on the German Jumo 004
turbojet engine in 1942. Later, although prior to the out-
WADC TN 55-298 33

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 CONFIDENTIAL
break of hostilities in Korea, a series of similar tests
were conducted on Americin engines of a later design in the
Power Plant Laboratory, Wright Field.

Many of the problems which had to be resolved during

PTI development and testing were similar to problems which
were encountered in previous afterburner developments.
These problems however were aggravated by the limitations
placed on the system in the form of weight and available
space restrictions. Since PTI was to be installed in the
F-86F aircraft, a rather large portion of the testing had
to be centered around obtaining adequate cooling of the
aircraft structure. Two test aircraft were used, one was
devoted to engine testing and the other toward resolving
the problems associated with the installations.

B. Testing

Thrust augmentation during PTI operation was obtained

by a combination of afterburning and basic engine overtem-
perature operation. Naturally basic engine overtemperature
was undesirable but it was nevertheless necessary because
of the space limitations in the aft fuselage of the F-86
airdraft. Had space been available, a larger tail-pipe and
nozzle could have been utilized, much like a conventional
afterburner, and the increase in back pressure on the en-
gine due to afterburning could have been reduced. It is
noted that afterburning has much the same effect as closing
the exhaust nozzle of an engine. A decrease in the exhaust
nozzle area increases the back pressure on the turbine
which reflects forward through the engine to the compressor
outlet, resulting in an increase in compressor pressure
ratio. In order to maintain the additional work from the
turbine that is required to maintain engine operation at
this increased compressor pressure ratio condition, the
turbine inlet temperature must be increased. Such was the
case with PTI where it was not possible to increase the ex-
haust nozzle area as much as was desired and still main-
tain stable burning. Without overtemperature, the engine
rpm would continually drop off with increasing altitude.
The degree of engine overtemperature for which the PTI
system was designed was based upon an estimated five hour
turbine bucket life. However, in order to bbtain stable
PTI burning, it was necessary to operate at augmentation
ratios higher than those originally intended with a conse-

WADC TN 55-298 34

CONFIDENTIAL
 CONFIDENTIAL
quent adverse effect on engine life. A variable-area exhaust
nozzle was used to maintain the correct combination of pres-
sure and temperature for stabilized PTI burning in the engine
exhaust section as well as to provide the correct nozzle area
for dry engine operation. In order to minimize control sys-
tem design complexities and to maintain the fuel flow require-
ments within the capacity of the existing fuel pumps, PTI
operation was limited to altitudes above 20,000 feet. Such
requirements which would allow operation below an altitude
of 20,000 feet were beyond the limits of the equipment used
in the PTI system.

The PTI system is shown schematically in Figure 24.

Fuel for pre-turbine injection was supplied from the normal
aircraft fuel system by the standard engine emergency fuel
pump. A PTI metering valve regulated the fuel flow in
accordance with the pressure schedule set by the EC-2 emer-
gency fuel regulator normally provided on the engine. The
metered PTI fuel was injected into the gas stream by means
of four probes located in alternate combustion chamber
transition liners forward of the turbine nozzle. The fuel
vaporized and burned downstream of the turbine wheel. Opera-
tion was such that flameholders were not thought necessary.
Naturally when the correct combination of pressure, tempera-
ture, and fuel-air ratio occurred, burning would take place.
The fuel in order to burn had to enter a zone that would
continually meet the combustion requirements. Thus the
conditions at turbine outlet had to meet such requirements.
Should combustion not have taken place, the pTI fuel
would have merely passed out through the engine unburned.
Once lit, the flame did not progress upstream as the gas
velocity far eýcceeded the rate of flame propagation. At
,first the system utilized a hot streak ignition system but
subsequent flight tests demonstrated that it was unnecessary
for satisfactory light offs. Initial flight testing was
accomplished using a water injection type tab to vary the
exhaust nozzle area pending development of a suitable
variable-area nozzle and nozzle control. Early tests were
run with the standard F-86F tail-pipe and it was not ade-
quate due to the high failure rate. Numerous tests, with
tail-pipes constructed of various materials and thicknesses,
were conducted. Since failures were caused by excessive
temperature of the tail-pipe skin, some method of lowering
the skin temperature was sought. The initial testing was
conducted with an insulation blanket surrounding the tail-
pipe. Radiation heat shields were added to several major

WADC TN 55-298 35

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 CONFIDENTIAL"

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T•. 55-298 Lwiw I

CONFIDENTIAL
 CONFIDENTIAL
aircraft frames and local cooling was provided for the aft
canted frame. Ram air scoops were incorporated to increase
the cooling airflow throughthe aft fuselage. High tail-
pipe skin temperatures measured during flights indicated
that a blanketed tail-pipe configuration was unsuitable
and that the temperature could be lowered by about 1500 to
2000F by their removal. The tailpipe blankets were thus
replaced by a silver plated shroud. The cooling problem
was made difficult by the partial loss of ejector action
due to the decrease in secondary air area as a result of
installing a variable-area nozzle which was necessarily
larger than the standard tail-pipe's nozzle. Further re-
duction in the temperatures to which the tail-pipe was
exposed became possible with the introduction of a corru-
gated louvered liner. The liner was installed for better
cooling and not to prevent screech since that problem was
not present. The liner was identical in construction, al-
though not in size, to the liner in the afterburner of the
J47D series engines used in the F-86D aircraft. Tail-pipe
construction was revised to include a flex joint between
the engine tail-cone and tail-pipe to eliminate the high
overhana moment of this tail-pipe configuration.
A by-pass fuel line with a variable trimmer valve was
specially provided between the engine small and large slot
manifolds to increase the basic engine fuel flow during
PTI operation. A solenoid shut-off opened the by-pass line
when PTI was selected. The variable trimmer valve is a
metering device which varies the by-pass flow in accordance
with small slot fuel pressure, closing completely above
approximately 45,000 feet altitude where the main engine
regulator fuel schedule is adequate to supply engine fuel
demands. A second special by-pass fuel line with a ground
adjustable trimmer (needle valve) was provided between the
small and large slot manifolds for ease in adjusting the
engine acceleration schedule on the ground.
After a review of many variable-area nozzle configura-
tions, the flat plate orifice type appeared to best suit
all needs since it was simple and relatively light in weight.
Activation was obtained at a single circumferential point
thus reducing the space requirements. Actuation loads were
relatively light since only friction had to be overcome.
The power source for the nozzle was a 1/6 horsepower air-

WADO TN 55-298 37

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 CONFIDENTIAL
frame mounted electric motor. The motor actuated a ball
bearing Jack screw mounted on the nozzle housing. Figure 25
is a sketch of the nozzle arrangement. It was found that
one nozzle position would suffice for PTI operation but two
were needed for dry operation; this was due to the fact
that the flow coefficient and effective areas varied appreci-
ably at the low actual values. There were two possible ways
of compensating for the variations of these factors -- con-
stantly variable or a step control nozzle. The simplest
method and also the most advantageous from a development
time standpoint was step control. Ideally, for the most
effectively controlled engine, a constantly variable-area
nozzle was desirable since any use of a step control would
necessarily compromise performance. It was decided that
only a fully variable-area nozzle would be satisfactory,
for with a step control a three position nozzle was neces-
sary; one for take-off, one for climb and normal performance
at altitude, and one for PTI operation.

The fully variable-area nozzle used in conjunction

with PTI was controlled by a pressure sensing device known
as the Micro-Jet. The unit controlled turbine discharge
temperature by varying the nozzle position to maintain a
constant pressure ratio across the turbine for any given
operating condition. The Micro-Jet contained a diaphragm,
one side of which was exposed to turbine discharge pres-
sure through a sensing line. The other side of the dia-
phragm was exposed to a pressure which was controlled by
bleeding compressor discharge air through a fixed inlet
orifice and a variable discharge orifice. The fixed ori-
fice was analogous to the engine's turbine and the orifice
size was ground adjustable to give the proper turbine
pressure ratio (determined by turbine discharge temperature).
The second orifice simulated the variable-area exhaust
nozzle. A tapered needle in the variable orifice was auto-
matically positioned by the diaphragm to relieve any pres-
sure differential across the diaphragm by increasing or
decreasing the pressure drop through the orifice. Should
the needle be disturbed from its neutral position (deter-
mined by the pre-set turbine pressure ratio) by a pressure
differential across the diaphragm, the electrical contacts
of the Micro-Jet would close signalling an electrical con-
trol box to energize the nozzle actuator motor to either
close or open the exhaust nozzle (depending on the direc-
tion of needle movement). The nozzle closed or opened

WADC TN 55-298 38

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7 -
'I CONFIDENTIAL

FIGURE 25. F-86F Tall-Pipe Variable Area Nozz.le

Utilized In The J47-GE-27 Engine Pre-
Turbine Fuel Injection Tests.

FIGURE 26. J47-GE-27 Engine Pre-Turbine Fuel

Injection Control Cockpit Presentation
In The F-86F Aircraft.

WADC TN 55-298 39

CONFIDENTIAL
 CONFIDENTIAL
until It reached a position which restored the turbine pres-
sure ratio to the value determined by the neutral position
of the needle.

A solenoid operated bleed valve on the compressor dis-

charge pressure side of the Micro-Jet diaphragm, which was
open during dry operation and closed during PTI operation,
permitted engine operation on two different turbine pres-
sure ratios thus allowing the control to be used for both
dry and PTI operation. Since retarding of the throttle be-
low its military power lowered the turbine pressure ratio,
the control tended to open the nozzle as the throttle set-
ting was reduced. The light-off problem at altitude was
considerably alleviated with the introduction of the Micro-
Jet control.
A PTI pressure out-out switch was provided as a means
of de-energizing the PTI circuit to prevent engine over-
speed should override of the main fuel control system by
the emergency fuel control system have occurred during PTI
operation. Figure 26 is a sketch of the PTI presentation
in the cockpit. PTI operation was initiated by the pilot
first setting the guarded PTI ready switch to ON, he then
set the nozzle selector switch to AUTOMATIC and then fol-
lowed by moving the throttle to the military power position
and then momentarily outboard from the military power de-
tent. PTI light-up and operation was fully automatic fol-
lowing the above three step procedure. However, the pilot
had to observe the PTI exhaust gas temperature limit of
1200 0 C during PTI operation. If the temperature approached
the limit, the throttle had to be retarded. PTI would re-
main in operation until the throttle was retarded to the
96% rpm position, but below 96% rpm PTI was automatically
shut off. If for any reason the pilot wished to stop PTI
operation, either the PTI ready switch could be positioned
to OFF or the throttle retarded past the 96% rpm position.

During 894 hours of static testing, 117 hours of PTI

were accomplished. In addition 46 hours of testing were
accomplished in an altitude tank with 8 hours of PTI being
accomplished. A total of 209 flights were conducted uti-
lizing two aircraft; during the 135 hours of flight testing,
17 hours of PTI were accomplished. A single 50 hour endur-

WADC TN 55-298 40

CONFIDENTIAL

S4.
 CONFIDENTIAL
ante test and three consecutive 50 hour tests were run.
During the single 50 hour endurance test 100 minutes of PTI
were accomplished. Five hours of PTI were accomplished
during the three consecutive 50 hour tests. The maximum
altitude reached during the testing was 53,760 feet. Twelve
flights were conducted above an altitude of 50,000 feet and
26 flights were conducted above an altitude of 45,000 feet.
C. Results

PTI applied to the J47-GE-27 engine in the F-86F air-

craft considerably improved the weapon's performance.
Pertinent data concerning PTI are shown in Figures 27
through 34. Also shown are performance data for a produc-
tion F-8G6F aircraft for purposes of comparison.
As can be seen from Figure 27, the rate-of-climb of
the F-86F aircraft using PTI, as compared to a standard
unaugmented F-86F, was more than doubled at an altitude of
35,000 feet, tripled at an altitude of 40,000 feet and was
increased by as much as four times at an altitude of 45,000
feet. Thus the time to climb from an altitude of 20,000
"* feet to an altitude of 45,000 feet was reduced by over six
minutes, the comparative times to climb being approximately
4 and 10 minutes for a PTI equipped aircraft and a produc-
tion aircraft respectifully. A PTI climb could be made
from an altitude of 20,000 feet to an altitude of 50,000
feet in approximately 6 minutes; Figure 28 presents such
data. The maximum level flight true air speed ofthe F-86F
aircraft was increased in the order of 20 to 25 knots be-
tween altitudes of 20,000 and 45,000 feet; Figure 29 shows
this increase. The thrust augmentation obtained by using
PTI, as can be seen from Figure 30, showed as average in-
crease of about 45% over the 20,000 to 45,000 feet altitude
range. As a result of PTI, a substantial improvement in
acceleration was made possible as evidenced from Figure 31.
Thus with PTI, the maximum level flight true air speed of
a production F-86F aircraft at an altitude of 35,000 feet
was reached over one minute sooner. Figure 32 is & com-
pilation of specific fuel consumption data; the Figure shows
the rather large increase in specific fuel consumption with
altitude when using PTI. As can be observed, the unaug-
mented J47-GE-27.engine's specific fuel consumption is near-
ly constant over the complete altitude range while the spe-

WADC TN 55-298 41

CONFIDENTIAL
 CONFIDENTIAL

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WAD- TN 55-298 44
.CONFIDENTIAL
 CONFIDENTIAL,
cific fuel consumption, utilizing PTI, increases due to the
necessity of having to increase quite substantially the
fuel flow at the higher altitudes. Nevertheless, the spe-
cific fuel consumptions obtained with PTI were considerably
better than those obtained with a standard afterburning en-
gine; for instance, the J47-GE-17 engine which powers the
F-86D aircraft has a specific fuel consumption approximately
40% greater than a PTI equipped J47-GE-27 engine. It was
concluded that the combat fuel requirements of PTI would
not impose a serious operational limitation; the combat radi-
us of the F-86F aircraft utilizing PTI was reduced even
less du6 to the addition of the increased weight of the
system. The J47-GE-17 engine utilizing a standard after-
burner gives a higher augmentation ratio than the PTI
equipped J47-GE-27, but also at a cost of nearly five times
the weight. With PTI, there was an increase in aircraft
weight of 140 pounds. The PTI kit itself actually weighed
210 pounds; however 70 pounds of existing parts were de-
leted. For instance, the tail-pipe and nozzle included in
the kit replaced those already in the aircraft. The after-
burner of the J47-GE-17 engine alone weighs 655 pounds.
A pronounced improvement in airplane maneuverability,
as can be observed from Figures 33 and 34 was possible
with the additional thrust provided by PTI operation. Al-
titude turns and maneuvers with PTI could be performed
with less drop-off in speed and altitude. Also, constant
altitude, constant speed maneuvers could be accomplished
with higher load factors and reduced turning radia with
FTI. One disturbing factor arose however due to the added
weight of the PTI installation. The addition of the PTI
installation plus. ballast, combined with the expenditure
of ammunition and fuel sequencing produced unacceptable
loading conditions. With the PTI system tested, 150 pounds
of ballast were required to provide the same acceptable
aircraft balance as an unmodified aircraft. The added
weight of the PTI installation caused a rearward shift of
the aircraft's center of gravity such that it exceeded
the aft neutral stability limits when the ammunition was
expended and thus the ballast in the nose of the aircraft
was necessary to correct that condition; but, when the
ballast was added and the ammunition was retained, the
center of gravity shifted forward to its maximum in-flight
position. The net result being an approximate 15% reduc-
WADC TN 55-298 45

CONFIDENTIAL

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 CONFIDENTIAL
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tion in the aircraft's maximum allowable load factor. It

was felt that with the addition of the proposed 12 inch wing
tip extension, the 150 pound ballast could be deleted. The
12 inch wing tip extension would allow an approximate 3% im-
provement in aircraft stability.

The degree of engine overtemperature for which the PTI

system was designed was based upon an estimated five hour
turbine bucket life. However, in order to obtain stable
PTI burning at the higher altitudes, it was necessary to
operate at augmentation ratios higher than those originally
intended with a consequent adverse effect on engine life.
Nearly 40% of the augmentation obtained was due to over-
temperature of the basic engine itself. Turbine buckets
during the test program had to be replaced nearly every one
hour of PTI operation. The buckets were replaced because
of excessive growth or actual failure. It was necessary
to replace the turbine wheel after approximately three hours
of PTI operation. Throughout the entire test program, five
turbine wheel failures were experienced and several others
were rejected after inspection. The exhaust-cone and tail-
pipe had to be replaced on the average of about once every
tenth flight. Some difficulty was experienced with the
variable-area nozzle as it had a tendency to stick; the
difficulty was overcome by hardening of the segments.
Burner roughness was found at altitud~es near 50,000 feet,
but in general the entire system functioned satisfactorily
and burning and light-ups were smooth. One very great
problem centered around the obtaining of the proper fuel
scheduling and many flights were ac'omplished to obtain
satisfactory operation, particularily at the higher alti-
tudes. Since the EC-2 standard emergency fuel regulator
was used as th6 PTI fuel metering device, maintaining a
close tolerance on the fuel injection rate was impossible
and that complicated with the narrow operating range of the
burner at high altitudes made the problem more difficult.
Although the emergency fuel regulator was part of the PTI
system, normal operation of the emergency fuel system was
possible.
A large portion of the test program was concerned with
determining an adequate configuration for the aft fuselage
since PTI operation necessarily overheated the aircraft's
WADC TN 55-298 47
CONFIDENTIAL
 CONFIDENTIAL

structure in that section; cooling air inlet and outlet

scoops were added to the aft canted frame in addition to
stiffener plates in two places on the inner side and straps
in three places on the outside. It was also necessary to
remove a lower portion of the aspiration as well as to add
heat reflecting shields. Access doors were placed to al-
low servicing and adjustment of the nozzle actuator. A
complete satisfactory configuration for the aft fuselage was
developed and a service bulletin was formulated. One of
the design objectives of the PTI systems was that it might
allow installation in the field; such an objective was re-
alized but approximately 600 man-hours would be necessary
for the installation and rework operation.
The severe reduction in parts life was obviously the
greatest deterrent to the use of PTI. It was estimated
that it would be possible to perform six missions, utiliz-
ing 6 to 10 minutes of PTI per mission, before it would be
necessary to replace turbine buckets. In addition, at
least one of the PTI missions would be necessary to check-
out and adjust the system; particularly critical was the
determining of the proper fuel schedule for operation above
an altitude of 45,000 feet. Near the end of the testing
program, some work was accomplished utilizing flameholding
elements in an effort to increase the life of the turbine
buckets which had previously served the purpose of flame-
holders.

WADC TN 55-298 48
CONFIDENTIAL
 CONFIDENTIAL
SECTION VI

GENERAL CONCLUSIONS

Without overtemperature, no increase in thrust is made

possible through overspeeding the rotor of the J47-GE-13
engine.

Imnediate thrust augmentation of the J47 engine may be

obtained through overtemperature operation. Such overtempera-
ture operation requires only the decreasing of the exhaust
nozzle area by placing restrictor segments in the tail-pipe.
This tabbing operation i easily accomplished and requires
only a small expenditure of work. Other than the decrease
in the life of the engine's hot parts subjected to overtem-
perature, engine operation is unaffected throughout the en-
tire operational range of the F-86 aircraft. The rate-of-
climb of the F-86E aircraft may be increased by as much as
four times at an altitude of 45,000 feet by overtemperaturing '
the J47-GE-13 engine by 18%. At lower altitudes, the increase
"in the rate-of-climb of the F-86 aircraft resulting from
overtemperature operation will not be as pronounced. Over-
temperature operation of the J47-GE-13 engine may be safely
utilized for a maximum period of ten minutes before a hot
parts inspection and turbine bucket replacement is necessary.
Overtemperature operation of the J47-GE-27 engine in the
F-86F aircraft is expected to offer a slightly lower increase
.in performance, relatively speaking when compared with the
J47-GE-13 engine in the F-86E aircraft due to the formers
initially better altitude characteristics; hot parts' life
would also tend to increase slightly as the J47-GE-27 engine
incorporates improvements in its hot section.

Liquid nitrogen injection into the J47 engine is not

practicle for use in the F-86 aircraft. Further work with
compressor refrigerant injection should be carried out on
advanced engines as it would undoubtedly prove worthwhile
for supersonic applications.

Many factors combine to make water-alcohol injection

into the J47 engine undesirable for use in the F-86 aircraft
although substantial increases in performance were demonstrated
in flight tests. The necessity of having a dual tank system

WADC TN 55-298 49

ý CONFIDENTIAL
° CONFIDENTIAL

and the lack of a~n adequate fluid injection metering control

would compromise the overall performance of the P-86 aircraft.
Without further development of the water-alcohol injection
system, unsatisfactory operation of the engine could be ex-
pected above an altitude of 45,,000 feet. Installation of a
water-alcohol injection system into the F-86 aircraft is a.
major task although no redesign of the aircraft structure is
necessary. The J47 engine may be expedited to hold up under
two hours of wsa-,.er-alcohol injection before an inspection is
necessary.

Pre-tutrbine fuel injection met all Initial performance

expectations with the exception of hot parts' life. Opera-
tion of the PTI system In the F-86 aircraft is satisfactory
up to an altitude of 45,000 feet and may be considered mar-
ginally satisfactory above that to altitudes slightly in ex-
cess of 50,000 feet. The rate-of-climb of the F-86F aircraft
is increased over four times at altitudes greater than 45,000
feet and the maneuverability and overall performance of' the
Aircraft is increased substantially above an altitude of
20,000 feet. 0vertemperature operation of the basic engine
during pre-turbine injection alone provided nearly 40% of
the total augmentation provided by PTI and although the PTI
portion of the system can have a reasonably lengthy parts!
life, the basic engine's hot parts would be limited to less
than one hour of PTI operation. The PTI system may be in-
stalled in the field as installation criterion is available
in service bulletin form, however, it might pose somewhat
of a task and installation had better be accomplished at a
depot. Fsrther work on a system employing the pre-turbine
injection principle is warranted although utilizing a more
modern engine than the J47. Such a system offers a means
for achieving a low dry loss, light weight afterburning system
which could possibly find application for take-off purposes
thereby not encountering the usual difficulty with after-
burning at altitude and also necessitate only a mlhimum of
complication through control requirements. Work should be
directed toward Improving the vaporization cooling of the
turbine through better fuel distribution and the investiga-
tion of flameholders, possibly of the retractable type.
Interstage turbine fuel injection on multistage engines should i
be further Investigated. Such work by the U.S. Navy, utiliz-..
Ing a J46 engine, haLs been relatively unsuccessful to date.

WADC TN 55-298 50

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 CONFIDENTIAL

Means of increasing the performance of the F-86 air-

craft were demonstrated through the program and such means
would have considerably enhanced the aircraft's combat cap-
ability as may be observed by superimposing the previous
discussed data upon similar data of the MIG-15 aircraft
given in the appendix for comparative purposes. However,
the most serious deterrent to the possible use of any of
the augmentation systems brought under development in the
program was the severe reduction in engine parts' life
which was brought about by their use. Although the J47
engine is basically a reliable, easily maintainable engine,
which can take considerable maltreatment, it is prone to
turbine wheel failures and any operation which subjects the
engine to overtemperature should only be resorted to in dire
emergencies. Attacking the problem at its source, by pro-
viding more durable basic engine hot parts, was beyond the
stcope of the program.
Each phase of the project was typlified by rapid pro-
gress in the beginning with further improvements requiring
intensified effort and considerable time as each system be-
came more complicated and increased altitude requirements
demanded more sophistibation. Refinement of each system
soon reached a saturation point where returns were dimin-
ishing due to the inherent limitations of the basic engine
and the high density of the aircraft's structure.

It is obvious that each of the thrust augmentation

means developed for the J47 engine In the F-86 aircraft was
of a strictly war emergency type and only suitable for use
in an extreme emergency and then only when adequate logistic
and maintenance support can be made available or where flight
safety or the aircraft in-commision rate can be compromised
for the increased performance gains.
The thrust augmentation program for the J47 engine in
the F-86 aircraft demonstrated that there is no substitute
for a basically better engine which Is made available through
a normal development program. This point was illustrated
in the superior performance and combat record of the F-86F
aircraft which uses the J47-GE-27 engine as compared to the
F86E which is powered by the J47-GE-13 engine. The J47-GE-27
engine is a development of the J47-GE-13 engine.

WADC TN 55-298 51
CONFIDENTIAL
 II

CONFIDENTIAL
REFERENCES

The project's official engineering file, which

was maintained by the Power Plant Laboratory, Wright
Air Development Center served as the primary reference
for WADC TN 55-298 in addition to information gathered
from personnel who participated in the program. Aside
from the project's record books, correspondence, minutes, and
memorandums, both Government' and Contractors' reports were
maintained in the file and the more significant rep-orts
which were used as references are given below.

Air Force Flight Test Center. Altitude Thrust Augmentation

Using Water-Alcohol Injection. AF Technical Report No.
A11°1•TU b6-8. March 1953. (Confidential Report)

Air Force Flight Test Center. Phase II Performance And

Serviceability Tests Of The F-86F Airplane USAF No.
5--13506 With Fre-Turbine Modifications. AF Technical
Report No. AFFTC 54-16. June 1954. (Confidential Report)

General Electric Company. Flight Test Of Overtemperature

War Emergency Power Setting On A J47-WE,13 Engine Installed
TnX' F-86e Aircraft At AFTC Edwards, California. GE
Bulletin No. DF52GT745. 23 June 1952. (Unclassified Report)

General Ele ctric Company. Estimated Minimum Performance

Of The General Electric J47-GE-27 Turbojet Engine At
War Emergency Rating With Preturbine Injection Kit.
GE Bulletin No. R53AGT568. 15 September 1953.
(Confidential Report)

General Electric Company. Model Specification War

Emergency Rating System Kit No. 7032R37. GE Specification
No,. E-641. 31 December 1953. kOonfiderti&l Specification)

General Electric Company. Pre-Turbine Fuel Injection for

Thrust Augmentation. GE Technical Information Series
No. R55AGT95. 11 March 1955. (ConfidentiAl Report)

North American Aviation, Inc. Thrust Augmentation Of A

J47-GE-1i Engine By Means Of Liquid Nitrogen Injection
Into The Compressor Inlet. NAA Report No. NA52-297.
i7 dune 1952. (onhridentlal Report)

North American Aviation, Inc. F-86F Airplane With Pre-

Turbine Injection (PTI) Thrust Augmentation. AJA Summary
Report No. NA-54-664. 9 Jul 1954. (Confidential Report)

VWADC TN 55-298 52

CONFIDENTIAL "
 CONF/DENTIAM
National Advirory Committee For Aeronautics. (Confidential
Title) Altitude Investigation Of Thrust Augmentation Of A
"J47-GE-27 Turbojet Engine By Injecting Additional Fuel
Immediately Ahead Of The Turbine. NACA Research Memorandum
RM E53L31, 3B December 1953. (Confidential Report)

Air Technical Intelligence. (Secret Title). Technical

Report TR-AC-27. 13 October 1953. (Secret Report)
As the title of this report is Secret it can not be
given here. The material taken from the above report
and contained in WADC TN-55-298 is as of the writing
of this TN classified Confidential.

WADC TN 55-298 53

CONFIDENTIAL I
 CONFIDENTIAL
APPENDJIX

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-VWADO TN 55-298 54

CONFIDENTIAL 1

5, .GLPH-10637
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Reproduced by
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NOTICE: WHEN GOVERNMENT OR 011HER DRAWINGS, SPECIFICATIONS OR OTHERL DAT.

ARE USED FOR ANY PURPOSE OTHER THAN IN CONNECTION WrT£ A DEFXNT1ELY REL iEl:
GOVERNMENT PROCUREMENT OPERATION, THE U. S. G0)VERNMENT THEREBY !NCURI
NO RESPONSIBILITY, NOR ANY OBLIGATION WHATSOEV1R; AND THE FACT THAT THE
GOVERNMENT MAY HAVE FORMULATED, FURNISHED, CIR P'4 ANY WAY SUPPIALD THE
SAID DRAWINGS, SPECIFICATIONS, OR OTHER DATA It. (/.1 -1 j.,E . ,
IMPLICATION OR OTHERWISE AS IN ANY MANNER LIC Z.8iNG THF HC%,Dra, OF, ANY 4y"; -
PERSON OR CORPORATION, OR CONVEYING ANY RIG]•r,T OR PERMISSION TO MIN-UFAt J14
USE OR SELL ANY PATENTED INVENTION THAT MAY 2 .ANY WAY BE RELATED '!iH:> ,

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 DEPARTMENT OF THE AIR FORCE
HEADQUARTERS AIR FORCE MATERIEL COMMAND
WRIGHT-PATTERSON AIR FORCE BASE OHIO

FEB t 9 2002

MEMORANDUM FOR DTIC/OCQ (ZENA ROGERS)

8725 JOHN J. KINGMAN ROAD, SUITE 0944
FORT BELVOIR VA 22060-6218

FROM: AFMC CSO/SCOC

4225 Logistics Avenue, Room S 132
Wright-Patterson AFB OH 45433-5714

SUBJECT: Technical Reports Cleared for Public Release

References: (a) HQ AFMC/PAX Memo, 26 Nov 01, Security and Policy Review,
AFMC 01-242 (Atch 1)

(b) HQ AFMC/PAX Memo, 19 Dec 01, Security and Policy Review,

AFMC 01-275 (Atch 2)

> (c) HQ AFMC/PAX Memo, 17 Jan 02, Security and Policy Review,
AFMC 02-005 (Atch 3)

1. Technical reports submitted in the attached references listed above are cleared for public
release in accordance with AFI 35-101, 26 Jul 01, PublicAffairs Policies and Procedures,
Chapter 15 (Cases AFMC 01-242, AFMC 01-275, & AFMC 02-005).

2. Please direct further questions to Lezora U. Nobles, AFMC CSO/SCOC, DSN 787-8583.

LEOrRA U. NOBLES
AFMC STINFO Assistant
Directorate of Communications and Information

Attachments:
1. HQ AFMC/PAX Memo, 26 Nov 01
2. HQ AFMC/PAX Memo, 19 Dec 01
3. HQ AFMC/PAX Memo, 17 Jan 02

cc:
HQ AFMC/HO (Dr. William Elliott)
 DEPARTMENT OF THE AIR FORCE
HEADQUARTERS AIR FORCE MATERIEL COMMAND
WRIGHT-PATTERSON AIR FORCE BASE OHIO

MEMORANDUM FOR HQ AFMC/HO

FROM: HQ AFMC/PAX

SUBJECT: Security and Policy Review, AFMC 02-005

1. The reports listed in your attached letter were submitted for security and policy review IAW
AFI 35-101, Chapter 15. They have been cleared for public release.

2. If you have any questions, please call me at 77828. Thanks.

EESJA. MRO
3ecurity and Policy Review
S~Office of Public Affairs
Attachment:
Your Ltr 14 January 2002
 14 January 2002

MEMORANDUM FOR: HQ AFMC/PAX

Attn: Jim Morrow

FROM: HQ AFMC/HO

SUBJECT: Releasability Reviews

1. Please conduct public releasability reviews for the following attached Defense
Technical Information Center (DTIC) reports:

a. Flight Test Programfor Model P-86 Airplane Class - Jet Propelled Fighter, 2
December 1946; DTIC No. AD-B804 069.

b. Physiological Recognition of Strain in Flying Personnel: Eosinopenia in F-86

Combat Operations,September 1953; DTIC No. AD- 020 375.

c. Phase IV Performance Test of the F-86F-40 Airplane Equipped with 6x3-inch

Leading Edge Slats and 12-inch Extensions on the Wing Tips, May 1956; DTIC
No. AD- 096 084.

d. F-86E Thrust Augmentation Evaluation, March 1957; DTIC No. AD- 118 703.

e. F-86E Thrust Augmentation Evaluation, Appendix IV, March 1957; DTIC No.
AD- 118 707.

f. A Means of Comparing Fighter Effectiveness in the Approach Phase, October

1949; DTIC No. AD- 223 596.

g. War Emergency Thrust Augmentation for the J47 Engine in the F-86 Aircraft,
August 1955; DTIC No. AD- 095 757.

h. OperationalSuitability Test of the F-86FAirplane, 4 May 1953; DTIC No. AD-

017 568.

i. Estimated Aerodynamic Characteristicsfor Design of the F-86E Airplane, 26

December 1950; DTIC No. AD- 069 271.

j. Combat Suitability Test of F-86F-2 Aircraft with T-160 Guns, August 1953; DTIC
No. AD- 019 725.
2. These attachments have been requested by Dr. Kenneth P. Werrell, a private
researcher.

3. The AFMC/HO point of contact for these reviews is Dr. William Elliott, who may be
reached at extension 77476.

kJOHN D. WEBER
Command Historian

10 Attachments:
a. DTIC No. AD-B804 069
b. DTIC No. AD- 020 375
c. DTIC No. AD- 096 084
d. DTIC No. AD- 118 703
e. DTIC No. AD- 118 707
f. DTIC No. AD- 223 596
g. DTIC No. AD- 095 757
h. DTIC No. AD- 017 568
i. DTIC No. AD- 069 271
j. DTIC No. AD- 019 725