The effect of exhaust nozzle restrictions on the performance of turbo-jet engines

Citation preview

THE EFFECT OF EXHAUST NOZZLE RESTRICTIONS ON THE PERFORMANCE OF TURBO-JET ENGINES

A Thesis Presented to the Faculty of the Department of Mechanical Engineering The University of Southern California

In Partial Fulfillment of the Requirements for the Degree Master of Science

by Robert L, Sohn December, 1950

UMI Number: EP60515

All rights reserved INFORMATION TO ALL USERS The quality of this reproduction is dependent upon the quality of the copy submitted. In the unlikely event that the author did not send a complete m anuscript and there are missing pages, these will be noted. Also, if material had to be removed, a note will indicate the deletion.

Dissertation Publishing

UMI EP60515 Published by ProQuest LLC (2014). Copyright in the Dissertation held by the Author. Microform Edition © ProQuest LLC. All rights reserved. This work is protected against unauthorized copying under Title 17, United States Code

uest ProQuest LLC. 789 East Eisenhower Parkway P.O. Box 1346 Ann Arbor, Ml 4 8 1 0 6 - 1346

Mis..

'51

‘ otoSi

This thesis, w ritten by .....

under the guidance of h te —.Facuity Com mittee, and approved by a ll its members, has been presented to and accepted by the Council on Graduate Study and Research in p a rtia l f u lfill­ ment of the requirements fo r the degree of

Date.

Faculty Committee

Chairman

TABLE OF CONTENTS Section I II

III

Page INTRODUCTION...............................

1

SUMMARY OF RESULTS..........................

3

Effect of Nozzle Restrictions on Engine Performance............................

4

Thrust output ..........................

4

Specific fuel consumption ..............

4

Optimum bend tab setting................

5

Effect of Nozzle Restrictions on airplane Performance ..........................

6

DISCUSSION.................................

Y

Theoretical Investigation ................

8

Test Equipment...........................

9

Description of engine and tailpipe restrictions.........................

9

Instrumentation.......................

10

Static thrust stand ..................

10

In-flight thrust measurement.........

10

General airplane instrumentation. . . .

11

Test Procedure...........................

11

Static ground tests ....................

12

In-flight tests ........................

12

Data obtained.........................

12

Method of Data Reduction and Presentation •

14

Thrust measurement theory ..............

14

iii Section

Page Calibration factor......................

16

Static thrust presentation..............

16

Airplane performance analysis ..........

17

Discussion of R e s u l t s ....................

18

Static thrust and specific fuel consumption

•

18

Calculated results....................

18

Actual results........................

18

In-flight performance checks• ..........

20

Performance checks. ..................

20

Variation of engine operating paramet­ ers with speed and altitude........

21

Agreement with results obtained by NACA.............................

22

.

Conclusions and Recommendations..........

23

IV

NOMENCLATURE................................

25

V

REFERENCES..................................

26

VI

CHARTS AND FIGURES..........................

27

APPENDIX..

38

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

INDEX OP CHARTS AND FIG-URES

Figure

Page

VI -1

Variation of Gross Thrust Output with Exhaust Gas Temperature at Full Engine Speed........................................ 27

2

Variation of Specific Fuel Consumption with Exhaust Gas Temperature at Full Engine Speed ..................... .. • • • 28

3

Effect of Exit Nozzle Area on Gross Thrust Output and Specific Fuel Consump­ tion at Full Engine S p e e d . ................... 29

4

Nozzle Coefficients and Optimum Bend ................... . . . . 30 Tab Settings

5

Effect of Exhaust Gas Temperature on Engine Performance in Flight ............. . 3 1

6

Rectangular Nozzle Bend Tab................... 32

7

Annular Nozzle Ring........................... 33

8

Static Engine Thrust Calibration • • • • • • •

9

Exhaust Gas Survey Rake. • • • • ........... . 3 5

10

Cooling Air Diagram for a Turbo-Jet Engine • . 3 6

11

Schematic Diagram of Turbo-Jet Engine Showing Station Location • • ............

34

..37

LIST OF TABLES

Table I II

Page Effect of Nozzle Restrictions on Airplane Performance................................... Test Program...............................

6 13

I

INTRODUCTION

Production turbo-jet engines of the same series often produce varying exhaust gas temperatures and thrust performance under identical operating conditions because of variations in internal efficiencies or dimensions*

For

this reason tailpipes are usually provided to the airframe manufacturer 1% to 2% oversize in nozzle area, resulting in below normal exhaust gas temperatures for the majority of engines*

However, in order to achieve the maximum

power output it is necessary to maintain the maximum per­ missible gas temperature since gross thrust is a direct function of the exhaust gas velocity, which in turn is governed by the gas temperature* One means of controlling the exhaust gas tempera­ ture of turbo-jet engines which have fixed-area exhaust nozzles is to vary the nozzle area by means of tabs or restrictors.

Restricting the nozzle area demands an in­

crease in gas velocity which the engine control system brings about by increasing the fuel flow; the increase can­ not be achieved by increasing engine speed since RPM is definitely limited by governor action.

Injecting addition­

al fuel will of course raise the turbine operating tempera­ tures, but not by a critical amount*

Turbine temperature

is in effect kept within limits by observing the maximum

allowable tailpipe temperature, which dictates the extent to which the nozzle area may be restricted*

Precursory

calculations showed that thrust gains of approximately 9% could be achieved for the average production jet engine of the 5000-lb class by properly regulating the nozzle area* It is the purpose of this paper to describe and summarize some of the tests which were conducted by North American Aviation in order to determine the actual thrust gains produced by, and the relative efficiencies of various types of tailpipe restrictors* Because of the nature of these tests a restricted classification has been imposed upon all data obtained thereof*

In order to conform with security regulations all

data are presented in percentage form and no actual figures are given.

However, it is not believed that this practice

has detracted from the factual worth of the paper but rather has presented the results in a more meaningful manner•

II

SUMMARY OF RESULTS

A theoretical investigation was undertaken to deter­ mine the performance gains that could be achieved by rais­ ing the exhaust gas temperature to the maximum allowable by means of nozzle area restrictions*

The calculations

showed that for the type of engine investigated 3*75 pounds of thrust increase per degree rise in exhaust gas tempera­ ture could be achieved.

Similarly, for the cruise condi­

tion the specific fuel consumption would be increased .00074 pounds of fuel per pound of thrust per degree rise in exhaust gas temperature. Complete ground and in-flight calibrations were then made of a turbo-jet engine with rectangular-type bend tabs installed in the exhaust nozzle as a means of controlling the exhaust gas temperature and thus thrust output.

The

clean tailpipe configuration was also tested, as was the tailpipe with annular rings installed.

The relative ef­

ficiency of the rectangular-type tabs at a given tailpipe temperature was determined by comparing the incremental performance gains achieved with the tabs with the gains achieved with the annular rings. Figures 1 through 5 of section VI, and Table I on page 6 summarize the tab effectiveness tests.

It is seen

that significant gains can be realized in the performance

of production turbo-jet engines*

Excellent agreement is

shown between the theoretical performance gains and those actually achieved with the annular restrictors. EFFECT OF NOZZLE RESTRICTIONS ON ENGINE PERFORMANCE Thrust Output; As shown in figure 1, at full RPM and at limiting tailpipe temperature the rectangular-type bend tabs in­ crease the static thrust output by 1*7 percent, whereas the annular rings increase the output by 9.1 percent under the same conditions•

In terms of exhaust gas temperature

the above results are .67 and 3.5 pounds of thrust increase per degree rise in gas temperature for the tabs and rings, respectively.

Assuming the annular rings to represent full

obtainable efficiency the rectangular bend tabs are 19 percent efficient. Specific Fuel Consumption: The specific fuel consumption of the engine with a clean tailpipe is adversely affected by the addition of rectangular tailpipe tabs (see figure 2)•

At limiting

tailpipe temperature and full engine speed the rectangular tabs increase the specific fuel consumption by 5.3 percent, whereas the annular rings preserve the specifics obtainable with an efficient restrictor.

5 Optimum Bend Tab Setting: The data presented in figure 4 show that the highest tab efficiency is reached at an approximate 5° setting, and decreases steadily as the tab angle is increased*

Pull

tailpipe temperature is reached with 35° setting for the particular engine tested*

It should be noted, however,

that gross thrust increases even though tab efficiency drops off, since the exhaust gas velocity is being increas­ ed. Tests conducted with another engine indicated that the maximum thrust increase could be achieved with a rec­ tangular tab setting of 65° (provided tailpipe temperature was not exceeded)• Setting the tabs to a higher angle re­ sulted in a thrust loss because of the greatly increased turbulence at the higher angle. The performance data were obtained with the engine installed in an airplane so that the effect of tabs on cooling air ejection would be included in the thrust values obtained.

Likewise, a complete flight test program at al­

titude was conducted in order to establish tab effective­ ness over a wide range of operating conditions.

The re­

sults obtained thereof were in good agreement with those derived from the ground tests (as shown in figure 5)• all cases the data were reduced to standard atmospheric

In

6

conditions and to a given altitude so that fair comparisons could be made. EFFECT OF NOZZLE RESTRICTIONS ON AIRPLANE PERFORMANCE In order to illustrate the effect of engine output (nozzle restrictions) on airplane performance a hypotheti­ cal fighter airplane is analyzed in the table below.

The

salient performance features are compared for various en­ gine outputs simulating an undersize, clean tailpipe, the tailpipe with rectangular bend tabs installed in the nozzle, and the basic tailpipe with an annular nozzle ring.

The

tab and ring configurations are set so as to produce limit­ ing tailpipe temperatures at full power static operation. TABLE I* EFFECT OF NOZZLE RESTRICTIONS ON AIRPLANE PERFORMANCE Rings Original Tabs Tailpipe Installed Installed 5000 5086 5455 Static Thrust (lbs) Rate of Climb at (ft/min) 5000 5132 5685 Sea Level 3000 4070 5770 25,000 ft Maximum Speed at (knots) Sea Level 503 500 500. 5' Take-off Distance (ft)

3000

2945

2725

Range at (naut. miles) 30,000 ft

1000

963

1014

^Basic data for the airplane performance were taken from reference (k)• The effects of the changes in engine out­ put due to the variation in nozzle area were incorporated into the performance by the methods given in the same reference•

Ill

DISCUSSION

As pointed out in the Introduction tailpipes for turbo-jet engines are supplied to the airframe manufacturer slightly oversize to reduce the temperatures of those en­ gines which would overheat with a normal-sized tailpipe. It is not practical to hand-select a tailpipe for each in­ dividual engine since an overhaul could change the internal characteristics of the engine sufficiently to render the original tailpipe useless; moreover, tailpipe temperature is dependent upon ambient air conditions, which in actual operation may vary widely from day to day* As a solution small rectangular bend tabs'"' or restrictors are usually inserted in the nozzle as a means of increasing the tailpipe temperature (which increases thrust by raising the exhaust gas velocity)*

These tabs

are easily bent on the ground to the desired setting, are light in weight, and are readily adaptable to any tailpipe. Although the bend tabs are desirable from a # It is well to define the terms, restrictors, tabs and annular rings as used in this paper. Restrictor is a general term used to indicate any device which restricts the tailpipe nozzle area. Tabs (or rectangular bend tabs) refer specifically to the restrictors illustrated in figure 6 and are the type employed in the effectiveness tests. The annular-ring type of restrictor is illustrated in figure 7 and is also the type used in the above tests.

8 fabrication and maintenance standpoint, they produce but a small percentage of the performance gains which are theore­ tically possible*

The test program described herein was

undertaken to determine the actual efficiency of the tabs* THEORETICAL INVESTIGATION A theoretical investigation was undertaken to deter­ mine the results that could be expected from the test pro­ gram*

The complete method used is presented in Appendix I

to this report along with a sample calculation to indicate the use of the equations which are derived therein*

The

approach taken closely follows the one described in refer­ ence (i)♦ No engine dimensions or RPM*s were required, only the conditions of the induction air*

Compression

ratios and turbine inlet total temperatures were taken from the available engine data, and compressor and turbine efficiencies of *80 and *90, respectively, were assumed* The efficiencies were held constant throughout the range of calculations* The thrust output and nozzle area were then computed i over a range of turbine inlet temperatures to match the ex­ haust gas temperatures obtained during the actual tests* Since the increase in total temperature across the combus­ tion chamber was known, the change in fuel-air ratio was computed and hence the variation in specific fuel

consumption* The results of this investigation are presented in figures 1 through 5 along with the corresponding test data for comparison purposes*

The close agreement between the

theoretical and test data will be noted* TEST EQUIPMENT Description of Engine and Tailpipe Tabs: The engine selected for these tests was a typical jet engine of the 5000-lb thrust class*

The engine is an

axial flow design utilizing a twelve-stage compressor, eight combustion chambers and a single-stage turbine; full operating speed is approximately 8000 RPM.

As installed

in the test vehicle, a typical present-day jet fighter, a ram nose duct supplies induction air to the engine, and a tailpipe 4*2 diameters in length carries the exhaust gases to the ambient air*

An ejector shroud surrounds the tail­

pipe exhaust nozzle, providing the pumping action for cool­ ing air over the tail cone and pipe* The tailpipe exhaust nozzle is approximately 20 inches in diameter at the exit point and is narrowed with o a 7 taper* The nozzle area of the production type nozzle is approximately 200 square inches* The following tabs and rings were tested: (1)

A rectangular bend tab (as shown in figure 6)

10 (2)

An annular ring (shown in figure 7)

The rectangular tabs were approximately 3 x 4

inches in

size and were easily bent to any desired angle of projec­ tion into the exhaust stream.

Two tabs were attached to

the nozzle exit, one at the top of the nozzle and the other diametrically opposite. Separate annular rings were fabricated for each of the nozzle configurations; the restrictors were theoretical­ ly as efficient as an integral tailpipe of corresponding area and thus could be used as a base for computing the relative efficiencies of the rectangular tabs. Instrumentation: The determination of usable engine thrust during these tests resolved itself into two phases:

first, mea­

surement of thrust during ground test runs, and secondly, measurement of thrust in flight. Ground calibrations of the engine as installed in the airplane were accomplished by attaching strain gages to linkages anchoring the landing gear struts during ground runs of the engine. in figure 8.

A diagram of this arrangement is given

An accurate calibration of the strain gages

was made prior to the tab tests and thrust measurements are accurate to within ± 1% . In order to measure the thrust output In flight a

11 survey rake was installed in the tailpipe to record the total pressure and total temperature of the exhaust gases. The sampling of these two properties of the exhaust gases enabled the thrust to be computed as described in a later section. The survey rake employed for the tests is depicted in figure 9.

Three total pressure tubes are equally spaced

on the left hand side of the rake and two shielded thermo­ couples are situated as shown on the right hand side.

The

total pressure tubes are connected to manifold pressure gages in a photo-recorder box in which time histories of the pressure readings are made when the pilot depresses a trigger switch on the control stick.

All other pertinent

data such as airplane speed and altitude, fuel remaining (for airplane gross weight evaluation and for specific fuel consumption), and engine RPM are recorded at the same time. The thermocouples are routed to a Brown recorder which like­ wise records time histories of exhaust gas temperature and outside air temperature.

The tailpipe temperature was re­

corded by the pilot whose indicator records the average of 4 thermocouple leads equally spaced around the inside per­ iphery of the tailpipe• TEST PROCEDURE The test program outlined for the tailpipe tab tests

12 consisted of two phases:

(l) a ground calibration of the

engine for each tab configuration installed, and (2) an in­ flight check of the results obtained during each ground run. Static Ground Runs: In all, seven ground runs were conducted*

The en­

gine was installed in the airplane throughout the test pro­ gram so that installation losses would be included in the final performance values.

In particular the effect of ex­

haust nozzle configuration on ejector pump losses could be noted (see figure 10)•

The table on the following page

lists the tailpipe restrictors tested* In-Flight Checks: Immediately following the ground run of each tab configuration a check calibration was made in flight* These flight tests determined the effect of widely varying engine operating conditions on tab efficiency.

Since the

basic drag of the airplane was accurately known from pre­ viously obtained data, airplane performance could be taken as a reliable indication of useful engine output*

Each

flight consisted of a take-off, climb to low cruise altitude at full RPM, a speed-power run at this altitude, a high speed level flight run at low altitude, and landing. Data Obtained: The engine data recorded during each run included:

Ill

DISCUSSION TABLE II

TEST PROGRAM TEST RUN 1

TAILPIPE

TAB OR RESTRICTOR

TEMPERATURE SETTING

Production-type

None

As obtained

2

t!

Annular Ring

Limiting Tailpipe Temperature Under Static Conditions

5

N

n

Limiting Tailpipe Temperature in Plight*

4

\\

tt

Intermediate Setting to Give Temperature Between That Obtained During Runs (1) and (2)

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

.

____________

________ _

W

5

_____________________________ ________

6 V

...

I

Bend Tabs _________

Setting to Give Temperature Equal to (2)

N

M

Setting to Give Temperature Equal to (3)

w

W

Setting to Give Temperature Equal to (4)

*Under ordinary conditions tailpipe temperature will decrease by approximately 60°P as the airplane is accelerated during take-off to flight speed. H

03

14 RPM, fuel flow, tailpipe temperature, exhaust gas tempera­ ture and total pressure.

Airplane data taken during flight

consisted of ambient air temperature, pressure altitude, pilot!s indicated airspeed, and time. METHOD OP DATA REDUCTION AND PRESENTATION Thrust Measurement Theory; The net thrust of a turbo-jet engine can be expres­ sed as the change in the momentum of the mass flow passing through the engine.

Algebraically, the net thrust is

equated to the gross or jet thrust output minus the ram drag, which is that component of the gross thrust required to accelerate the induction air to airplane velocity.

It

is the resultant force that may be utilized to overcome the aerodynamic drag forces acting on the airplane since all factors that influence power plant output are inherent­ ly included in the net thrust tern. The equation is: %

= pg - pR

The ram drag term, Fj^, is equal to the gross thrust, Fg, times the ratio of airplane velocity to exhaust gas velo­ city: t?_

R

s

tt

vairplane g Vvexhaust k --gas

The assumption made here is that the mass of induc­ tion air is equal to the mass of the exhaust gases.

15 Actually, fuel is added to the flow, approximately one part in sixty, hut this generally balances air hied from the compressor for cahin ventilation and cooling, and air lost from the compressor through leakage. By assuming the total pressure and temperature of the gas at the nozzle exit to he equal to that at the rake station (which is approximately 1.0 diameters upstream), the gross thrust can he evaluated since the static pressure is very nearly equal to ambient air pressure and the nozzle area is measured directly.

The complete equation combining

the variables mentioned above is:

F „ - ^ f Pk A [ ( W XfL- l ] ( i - ^ ) + A(R,-P„) where A

- nozzle area

P0 = ambient air pressure Pjj « nozzle pressure'* gas total pressure at rake station Vo = true airplane velocity Vj = exhaust gas velocy = ratio of specific heats of exhaust gas, equal to 1.335 for turbo-jet exhaust -*For choking conditions is taken as .54 Ptc-* last term of the above equation is an expression ror pressure thrust to allow for the expansion of the gases to atmos­ pheric pressure. For subsonic operation a P0.

16 References (c), (e) and (f) give a complete discus­ sion of the above method as well as the derivation of the flow equations involved. Calibration Factor; It is apparent that only the ideal gross thrust can be determined from the above equation.

The various assump­

tions of a non-viscous, non-turbulent flow, and discrep­ ancies in pressure and velocity integration will render the calculated thrust values inaccurate unless modified by an appropriate calibration factor.

This factor is obtained

by ratioing the actual thrust of the engine (as installed in the airplane) during a ground run to the calculated thrust computed for the same run.

The nozzle coefficient

is plotted as a function of the ratio of ambient air pres­ sure to exhaust gas total pressure. Static Thrust Presentation: Engine thrust data obtained during static ground tests is presented as corrected gross thrust versus cor­ rected RPM and tailpipe temperature.

These wcorrected”

values are defined as follows: 55 observed gross thrust x 3_ea level air pressure. ambient air pressure

17 A dimensional analysis of turbo-jet engine operation (references c and d) shows that curves presented in terms of the above parameters are essentially independent of atmospheric conditions.

Sea level standard day performance

can be read directly from the faired curve of corrected values• Likewise, fuel consumption data is presented as a function of corrected gross thrust in the following form: w* T 'lS

s observed fuel flow x

,.3£-Y.e-3- a,tr.,jpre.s^ar3 . ambient air pressure

v /sea level air temp\ J \ambient air temp J Airplane Performance Analysis: Airplane performance was reduced to NACA standard atmospheric conditions and to a performance gross weight by the methods set forth in references (f) and (g)•

A

direct comparison of airplane performance and, particularly, thrust available with the various restrictors was then possible.

Correction of thrust was accomplished by a

process similar to that employed for static thrust data with the addition of a step to allow for changes in free stream Mach number. Airplane range is given in terms of miles per pound of fuel as a function of Mach number for a given ambient air pressure to gross weight ratio.

The maximum range

values indicated by these curves for the various

18 restrictors are then readily determined* DISCUSSION OP RESULTS Static Thrust and SFC: Calculations were carried out using theoretical engine efficiencies and compressor ratios to indicate the performance gains which could be achieved by increasing exhaust gas temperature to the limiting value through changes in the exhaust nozzle area* ficiency was assumed*

No change in nozzle ef­

These calculations were verified

with excellent agreement (as shown in figures 1 through 5) by results obtained with annular restrictors installed in the tailpipe*

The anticipated pounds of thrust increase

per degree rise in tailpipe temperature was likewise fully realized in the test calibrations. The methods used in accomplishing these calculations were taken from references (i) and (h) with the pertinent equations summarized in Appendix I of this report. The actual effect of tailpipe tabs and ring restric­ tors on engine performance is clearly summarized in figures 1 through 5:

some output gains can be effected through the

use of tabs but much larger gains are realized by the use of annular rings*

Bend tabs are definitely detrimental to

SPC whereas the lov* specifics achieved with the clean tail­ pipe nozzle are essentially unaltered as the exit area is

19 decreased by the rings. An excellent indication of tab efficency is given in figure 4 in which is presented the ratio of measured to calculated thrust for various power settings, The primary reason for the low efficiency of the bend tab is the turbulent manner in which the exhaust gases are expelled in the region of the tab#

The resulting

nozzle losses decrease the gross thrust output of this type of engine by 7.4 percent.

Higher bend tab angles of course

mean greater turbulence and lower efficiency until finally a point is reached where steeper angles result in less thrust output.

With this particular engine the optimum

angle setting for the greatest thrust is approximately 65°. Since the primary purpose of this test was to de­ termine the effect of tabs with all installation losses included the tests were conducted with the engine in the airplane.

It becomes apparent after examining the inter­

nal airflow system of the airplane, figure 10, that nozzle configuration will directly influence ejector pump action. In general, the more directly the exhaust gases impinge upon the cooling air exit, and the greater the ejector out­ let overhang, the greater the internal drag losses of the installation.

Since it was considered possible that this

effect mig£it be contributing to the lower efficiency of the

20 bend tabs a test was conducted to isolate these losses, if any*

This was accomplished by disengaging the aft section

of the fuselage and noting the thrust output without ejec­ tor pump losses*

Within the accuracy of thrust measurement

cooling air losses were concluded to be of negligible pro­ portions • As was expected bend tab inefficiencies were reflec­ ted in SFC, which is presented in pounds of fuel per pound of thrust per hour*

The specifics were increased by 5*4

percent at limiting tailpipe temperature over the annular ring configuration.

The overall effect on airplane range

is described in the following section* In-Flight Performance Checks: In general the effect of tabs on engine performance was verified by the performance of the airplane with the various tab configurations installed.

Both the rates of

climb at full RPM and the cruise performance indicated the overall superiority of the annular rings.

The specifics

were increased 4.5 percent over the optimum for the cruise condition* The high speed checks were somewhat inconsistent* This is attributed to the power required characteristics of the airplane in this speed range, i.e., large changes in thrust produce small changes in speed because of the

21 sharply rising drag curve of the airplane.

The rate of

climb data, however, are considered to be very reliable since these data are functions of excess thrust and are particularly sensitive to small changes in power available* The fact that the in-flight checks verified the ground calibrations indicates that operation at sonic tail­ pipe velocities has no noticeable adverse effects on nozzle efficiency.

During the static tests at zero ram ttie ex­

haust gases were expelled at higji subsonic speeds, but as the airplane was accelerated to its maximum speed the high ram conditions caused the jet velocities to reach sonic speeds.

The total pressure into the tailpipe is increased

up to 25 percent at full ram which would probably tend to raise nozzle losses*

The increase in the maximum speed of

the airplane, however, corresponded roughly to the antici­ pated increase based on the static tests* Another conclusion that can be reached from the flight tests is that ejector pump action does not intro­ duce appreciable internal cooling losses even though the ejector is operating at greatly reduced back pressures* An interesting insight into the variation of the exhaust gas temperature over the operating range of the airplane was also gained from the flight tests*

It was

found that if the tailpipe temperature was limiting under

22 static conditions at full RPM, a 60° to 70° P drop was ex­ perienced as the airplane was accelerated during take-off to flying speed#

The temperature then remained fairly

constant at all altitudes and speeds until the airplane was climbed to extreme altitude, when the temperature would go overboard (probably due to lack of cooling)• This would indicate that additional performance gains could be achieved by installing a pilot-controllable nozzle tab, provided the tab were of an efficient design# Comparison of Results with NACA Data: References (a) and (b) describe tests conducted by the NACA to determine the efficiency of variable-area exhaust nozzles, as compared to fixed-area nozzles, for turbo-jet engines#

Included in the summary curves of these

reports were variations of thrust output versus exit area and versus exhaust gas temperature#

The data contained

therein are in almost exact agreement with the calculations and annular ring data taken from the North American tests. An approximate 4 pounds of thrust increase per degree rise in exhaust gas temperature was noted# As the NACA tests were conducted in the altitude wind tunnel of Cleveland engine performance data were ob­ tained over a full range of altitudes and free stream Mach numbers#

It is significant to note that the 4 pound per

25 degree increment given above for the static condition is applicable over the entire operating range of the airplane if it is modified by the ratio of air-flows between the flight and static condition. CONCLUSIONS AND RECOMMENDATIONS The conclusions drawn from these tests are: 1.

Static thrust gains of 9 percent can be achieved

with efficient tailpipe nozzle restrictors for the average production turbo-jet engine of the 5000-lb class. 2.

Rectangular-type bend tabs realize only one-

fifth of the maximum increase because of the greater tur­ bulence and consequent lower efficiency of these tabs. 3*

Rectangular-type tabs have an adverse effect on

specific fuel consumption whereas an efficient tab preserves the low specifics of the clean tailpipe. 4.

Static performance gains can be realized in

flight with no noticeable decrease in nozzle efficiency or increase in cooling air losses. 5.

The theoretical calculations of the effect of

area changes on the performance characteristics of turbo­ jet engines are in close agreement with the actual results* Recommendations are: 1.

Circular arc tabs of approximately 60° each

should be installed in the tailpipe nozzle as a means of

24 controlling the exit area and hence the exhaust gas velo­ city.

This type of tab would approach the efficiency of

the annular rings and yet could be easily installed in the number required for proper adjustment. 2.

Bend tabs which the pilot could control in

flight should not be installed since the small performance gains that could be achieved would not warrant the added weigjht and equipment involved (the rectangular-type tab would be the only practical type to install from a struc­ tural standpoint)•

IV

NOMENCLATURE Units

Symbols

sq ft

Nozzle exit area Specific heat at constant pressure

BTU/unit mass/°P

Internal energy

BTU/unit mass lbs

Thrust

BTU/unit mass

Enthalpy

slugs

Mass Mach number

in-Hg

Pressure

BTU/unit mass

Heat energy Temperature

°R

Thrust horsepower

hp cu ft/lb

Specific volume

ft/second

Velocity

BTU/second

Power Ratio of specific heats Efficiency Dens ity

slugs/cu ft Subs crfpts

C ompr es sor

j

jet

ambient, free stream

N

nozzle

total, stagnation

V

REFERENCES

NACA RMR E5H16, ”Test of an Adjus table-Area Exhaust Nozzle for Jet-Propuls ion Engines” , E. C. Wilcox, August 1945. NACA RM E8J25d, uInvestigation of Performance of Turbo-jet Engine v/ith Constant- and Variable-Area Exhaust Nozzles,” L. E. Wallner, November, 1948* ^Flight Testing,” Benson Hamlin, 1946. Bell Aircraft Aerodynamics Research Note 11, ”Applica­ tion of Dimensional Analysis to Jet Engines and JetPropelled Airplanes,” P. A. Lagerstrom, March, 1945. General Electric Report Df 81518, ttFlight Test Data Reduction Methods for Jet Power Plant Installations.” North American Aviation Report NA-47-1033, ”Methods of Flight Test Performance Data Reduction for Turbine Jet Propelled Airplane,” Revised October, 1948. British Report TN No. Aero 1348, ”Note on Various Methods of Performance Reduction for Jet-Propelled Aircraft,” F. Smith, December, 1943. California Institute of Technology Report No. 9230, ”Jet Propulsion,” published for the Air Technical Service Command, 1946* NACA ARRE6E14, ”Performance Charts for a Turbo-jet System,” Benjamin Pinkel and Irving M. Karp, June, 1946. NAA Report NA-46-1091, ”Flight Test Thrust Evalua­ tion of a Turbine-Jet Propelled Airplane in Flight,” September, 1947. ”Airplane Performance, Stability and Control,” Courtland D. Perkins and Robert E. Hage, 1949.

VARIATION OF GROSS THRUST OUTPUT WITH EXHAUST GAS TEMPERATURE AT FULL ENGINE SPEED 112

Increase in Gross Thrust — Percent

110 NOTE: 100 percent values represent clean tailpipe configuration

108 Calculated

106 Limiting Tailpipe Temperature

Annular rings 10i*

Rectangular bend tabs 102

—y 102 106 io h Increase in Exhaust Gas Temperature — Percent

Figure 1

VARIATION OF SPECIFIC FUEL CONSUMPTION V7ITH EXHAUST GAS TEMPERATURE AT FULL ENGINE SPEED Figure

11C

Note: 100 percent values represent clean tailpipe configuration Rectangular bend tabs

g

106

-P

Annular rings 102 Calculated

102

10li

106

Exhaust Gas Temperature — Percent

108

EFFECT OF EXIT NOZZLE AREA ON GROSS THRUST OUTPUT AND SPECIFIC FUEL CCMSUMPTION AT FULL ENGINE SPEED

29

Figure 3 Codes Calculated ----- Annular rings ----- Rectangular bend tabs Gross Thrust

115

Gross Thrust — Percent

110 Note: 100 percent values re­ present clean tailpipe configuration

105

•y —

x.00 Nozzle Area

Percent

Specific Fuel Consumption Specific Fuel Consumption — Percent

11