A performance analysis of a small turbojet engine at static sea level conditions

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A PERFORMANCE ANALYSIS OF A SMALL TURBOJET ENGINE AT STATIC SEA LEVEL CONDITIONS

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

In Partial Fulfillment of the Requirements for the Degree Master of Science in Aeronautics and Guided Missiles

■toy Harold M* Crawford

John T* 0*Keefe

UMI Number: EP60494

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T his thesis, w ritten by

Harold M. Cranford Vsilliam II, Stowell

John T. c*Keefe l/ilHam H, Woodward

under the guidance of fc.h3-2.3Facuity Com m ittee, and app ro ved by a l l its members, has been presented to and accepted by the C o uncil on G ra duate Study and Research in p a r t ia l f u l f i l l ­ ment of the requirements f o r the degree of

Master of Science in Aeronautics aM....Sulded...Mlssll.e.s.. une 1950

Faculty Committee

W & L ... Chairman

TABLE OP CONTENTS CHAPTER I.

II•

PAGE THE PROBLEM AND DEFINITIONS OF SYMBOLS USED

1

The p r o b l e m ................................

1

Statement of the p r o b l e m .................

1

Importance of the s t u d y .................

1

Definitions of symbols .....................

5

Organization of t h e s i s ................

7

REVIEW OF RELATED SUBJECTS ..................

9

Historical ..................................

9

Review of similar p r o j e c t s ........ ..

9

. • •

The University of Washington . . • • • • •

10

The Northrop Aeronautical Institute

10

The Cal-Aerotechnical Institute

III*

.

• • •

........

11

The Marquardt Aircraft Company • • • • « •

12

DESIGN, INSTRUMENTATION AND TEST PROCEDURE . . Design and construction Turbosupercharger

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

13

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

l6

Intake cone, ducting, and exhaust stack

•

Combustion c h a m b e r s ................... Starting system Fuel system Ignition system

13

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

l6 20

.

23 23 23

iii CHAPTER

PAGE I n s t r u m e n t a t i o n .............................

26

P r e s s u r e s .................................

26

Temperatures ...............................

28

Other I n s t r u m e n t a t i o n ....................

28

Test p r o c e d u r e ...............................

33

IV.REDUCTION OF EXPERIMENTAL R E S U L T S ..............

35

V.ANALYSIS OF P E R F O R M A N C E .........................

\\Z

Present performance

•

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

l\2

Comparison to an idealt u r b o j e t ..........

lf.2

Explanation of graphical presentation

[|_5

• •

Conclusions relative to present p e r f o r m a n c e .............................

56

Discussion of low p e r f o r m a n c e .............

5&

Low operating speeds ......................

56

Exhaust nozzle

56

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

Compressor and t u r b i n e ............... .. VI.SUMMARY, CONCLUSIONS, AND RECOMMENDATIONS

.

...

60 62

S u m m a r y ......................................

62

C o n c l u s i o n s .................................

63

R e c o m m e n d a t i o n s .............................

63

Exhaust n o z z l e .............................

63

Combustion c h a m b e r s ......................

6I4.

B I B L I O G R A P H Y ..........................................

65

LIST OP FIGURES FIGURE

PAGE

1*

Schematic Diagram of a Turbojet Engine

2*

Complete Turbojet Engine in Operating Position

15

3*

Cutaway View of Turbosupercharger.............

17

Side View of T u r b o s u p e r c h a r g e r ............. •

17

5»

• • • •

lij.

Bottom View of Turbosupercharger Showing Turbine Bucket Wheel

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

18

6 * Top View of Turbosupercharger Showing Compressor I n l e t ........................... 7«

18

Schematic Diagram of Engine Showing the Ducting

19

8 * Exhaust S t a c k ...................................

21

9*

22

Combustion Chamber from I~l6 Turbojet Engine

•

10*

Schematic Drawing of Fuel System

• • • • • • *

2I4.

11.

Schematic Drawing of Ignition System

.........

25

12*

Schematic Drawing of Instrumentation

.........

27

13*

Types of Instruments U s e d ......................

29

1lj_* Instrument Installation in Stagnation Tank End Plate 15*

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

30

Instrumentation at Station 5A in Exhaust Stack

30

l6 • Instruments in Ducting to Turbine Inlet • • •

•

31

17*

Instruments in Ducting to Turbine Waste Gate

•

31

18*

Thermocouple Components • • • . . • • • • • • •

IO.

rir’Qnh of*

TTCI DDM

32 11A

FIGURE

PAGE

20.

Graph of Temperatures VS R P M .................

I4.8

21.

Graph of Pressures VS R P M .....................

i^-9

22.

Graph of W a and (HP)t VS R P M ..................

51

23.

Graph of F and W f / W a VS R P M ....................

52

2lj..

Graph of C VS F ................................

53

25.

Graph of F AND RPM VS W f/ W a ....................

Sk

26 .

Graph of (Tg^/Tg0 )> (Ps^/^30 ^ (Psi/^)*^'

27*

* ^

*

VS R P M ..........................

Graph of Actual F and

Ideal F VS P3^/PSo

...

55 57

TABLE !• II•

PAGE Tabulated Test Data • « • Results of Computations

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

. . ,

37 ILL

CHAPTER I THE PROBLEM AND DEFINITIONS OF SYMBOLS USED The jet-propulsion system consisting of a compressor, a combustion chamber, a turbine, and a discharge nozzle, generally known as a turbojet, was extensively studied dur­ ing the war.

Since it has a relatively low engine weight

per pound of thrust developed and is able to provide thrust at high flight speeds it is an excellent power plant for military aircraft*

As far back as 19Mi small turbojet en­

gines using turbosuperchargers for the principal components were used to test various features of the engines and to familiarize students with turbojet operation.^

This chapter

presents the problem undertaken, the importance of the study definitions of the symbols used, and a discussion of the organization of the remainder of the thesis* I.

THE PROBLEM

Statement of the problem.

A group of seven graduate

students undertook the design, construction, instrumentation and performance analysis at static sea level conditions of a small turbojet engine.

The entire group participated in all

phases of the project, but the recording of the study was

Life Magazine. November 27. lQkk*

2 divided* report*

The performance analysis is the subject of this The design, construction, and instrumentation were

considered in another report#The conditions imposed on the design and operation were: 1#

A General Electric type B-31 turbosupercharger

and I-l6 combustion chambers were to be used as the principal components•

2#

The engine was to be operated at a speed less thai

twenty thousand rpm and at a nozzle box temperature below li^OQ°F for safety considerations# 3#

The engine was to be mounted vertically with no

attempt to construct an exhaust nozzle# Importance of the study#

The most important reason

for the conversion of turbosuperchargers to turbojets is to provide a simple turbojet engine with which studies relative to such engines can be conducted#

Some designers, however,

contemplate the use of such conversions as power plants for light planes or target drones# The total energy of the air which enters the thermo­ dynamic

system of the turbojet is increased by the additic

^ N. W# Tobey, A* R# Roth, and S# E# Salter, "Design, Construction, and Instrumentation of a Small Turbojet Engine (unpublished Master*s Thesis, University of Southern Calif-

of heat in the combustion chambers*

This additional energy

operates the turbine-compressor combination and greatly in­ creases the momentum of the gas as it passes through the engine*

This increase in the momentum of the gas leaving

the engine over that of the entering air appears as the thrust developed by the engine* The specific fuel consumption, expressed in pounds of fuel per

hour

per pound of thrust developed, m a y be con­

sidered as a measure of the overall efficiency of the engine a low specific fuel consumption being associated with a high efficiency* An investigation into the factors which result in a low specific fuel consumption consists primarily of evalua­ ting the efficiencies of the various engine components*

The

requirements of the various components for maximum engine efficiency necessarily result in design compromises. example, consider the compressor.

For

Engine thrust first in­

creases as the compression ratio increases due to the resulting increase in the cycle expansion ratio* beyond a certain optimum ratio,

However,

the thrust decreases.

This

condition is brought about by the limitation placed on the combustion chamber temperature by the blading of the turbine As the compressor ratio is increased more energy is required from the exhaust gases to drive the compressor*

Thus the

4 Moreover, as the compressor ratio is increased, the tempera** ture of the air undergoing compression increases*

Hence less

fuel can be burned in the combustion chamber before the max­ imum turbine inlet temperature is reached* is a decrease in thrust*

The net result

A combustion chamber designed for

maximum efficiency at a given fuel-air ratio can be operated at this ratio only so long as the temperature rise across the compressor plus the temperature rise due to combustion does not exceed the maximum allowable turbine inlet tempera­ ture*

Consequently,

too high a temperature ratio across the

compressor necessitates lowering the fuel-air ratio and results in lowering the combustion efficiency* Experience has shown that a jet propulsion engine will deliver approximately 50 lb* of thrust for every pound of air flowing through the engine per second*

Furthermore,

a modern rotary compressor requires about 100 hp* for each pound of air delivered per

second*3

It follows that, in an

engine developing 100 lb* of thrust, the turbine driving the compressor must produce about 200 hp*

"A reduction in com­

pressor efficiency of 1 per cent may require a 3 P©3? cent or greater increase in fuel to maintain constant power output.'*^3 Gr* Geoffrey Smith, M* B* E*, Gas Turbines and Jet Propulsion for Aircraft, (370 Lexington Ave., New York 17: Aircraft Books Inc., 19q.6)# ^ F* W. Godsey Jr. and Lloyd A* Young, Gas Turbines

5 Obviously, a compressor of low efficiency limits the power output of an engine and has a serious effect upon the over­ all efficiency* Prom the foregoing discussion it is clearly seen tha the specific fuel consumption and the overall efficiency of a turbojet engine are directly related to the efficiencies of the components of the system*

Since the component ef­

ficiencies are in turn effected by the compromises which must be made due to design limitations, an analysis of an engine1s performance is justified to permit an evaluation of the compromises which were made in its design and construetioi II.

DEFINITIONS OP SYMBOLS

A

Gross section area

Btu

British Thermal Unit, heat

ft^ to

raise 1 lb water 1°R c

Specific fuel consumption

Cp

Specific heat of air at constant pressure

cy

lb/hr/lb thrust

Btu/lb/°R

Specific heat of air atconstant volume

Btu/lb/°R

d

Diameter

ft

f

Friction factor

dimensionless

P

Thrus t

lb

6 h

Higher heating value of fuel

Btu/lb

H Pt

Rate of turbine input energy

HP

L

Length

ft

ra

Mass

slg

M

Mach number

Dimens i onle s s

P

Static pressure,

absolute

lb/ft2

(tabulated in inches of mercury) Stagnation pressure,

Ps

absolute

lb/ft2

(tabulated in inches of mercury) pn

Pressure in fuel line, guage

lb/in2

R

Gas constant for air

ft lb/sig °r

Re

Reynolds number

dimensionless

rpm

Rotational speed

rev/min

T

Static temperature

°R

T xs

Stagnation temperature

°R

V

Velocity

ft/sec

Wv',a

Weight flow of air

lb/min

Wf

Weight flow of fuel

lb/min

x

Ratio of specific heats, Cp/cv

dimensionless

Xo

Adiabatic efficiency of compressor 9 ratio of isentropic to actual temperature rise for a given pressure rise

X

CC

Combustion efficiency,

dimensi onle s s ratio of

cE

Energy efficiency, ratio of energy added to flow by compressor to dime ns i onless

that extracted by turbine ^ t

Adiabatic efficiency of turbine, ratio of actual to isentropic temperature drop for a given dimens i onless

pressure drop Viscosity

slg/ft sec

Density

slg/ft3

The following appearing as subscripts refer to location: o, 1 , 2 , 3 , ip, 5 , and 6

see Figure

c

compressor

average of values at

2 and 3

cc

combustion

average of values at

3 &nd 4

t

turbine

chamber

1 (page lip)

average of values at ip and 5 III.

ORGANIZATION OF THESIS

Since several similar projects have been studied their results were briefly discussed.

Northrop Aeronautical

Institute extensively uses projects on the conversion of turbosuperchargers as a means of instruction.

The University

of Washington and the California Aeronautical Institute have studied various components by use of such a conversion.

Mr.

West, at Marquardt Aircraft Company, has studied the problem

The design, instrumentation,

and test procedure are

discussed briefly since a more detailed discussion is available#^

The turbosupercharger and other components comprising

the actual motor are discussed first, followed by a discus­ sion of the fuel, ignition,

and starting systems#

The in­

strumentation is discussed in somewhat more detail to pre­ sent a sound basis for the conclusions drawn from these data# Test procedure is considered in detail to give an indication of precautions necessary for accurate test runs# The reduction of experimental data was conducted by the most accurate means available#

A typical reduction based

on a sample problem is presented in Chapter IV#

Since the

performance is below that expected of a well-designed turbo jet (because no exhaust nozzle was used) the analysis of per­ formance was divided into two portions.

The first considers

the actual performance as compared to that of an ideal turbo­ jet and presents, graphically, various relations relative to the actual performance.

The second portion discusses reasons

for the low performance and means of increasing performance# It is recommended that further experiments be conducted using a variable exhaust nozzle.

This testing could include

general performance tests through the entire range of opera­ tion or tests could be made on turbojet components or fuels#

CHAPTER II REVIEW OF RELATED SUBJECTS I.

HISTORICAL

In January 1930 Air Commander Frank Whittle of the Royal Air Force applied for his first patent on an engine design employing a gas turbine as the basis for jet propul­ sion.

The first flights by aircraft employing turbojet

engines as the sole means of propulsion were made in England and Germany in 19ljl.

With the appearance of the German high

performance turbojet fighter aircraft, the Me 262, as an operational airplane in World War II, interest in the turboje engine became widespread.

Due to its relatively low engine

weight per pound of thrust developed and its ability to pro­ vide propulsion at high flight speeds this type of engine has become extremely important as a power plant for military aircraft.

Its possibilities as a power plant for commercial

aircraft are being thoroughly investigated. II.

REVIEW OF SIMILAR PROJECTS

The remarkable performance obtainable from aircraft employing turbojet engines as a means of propulsion has attracted numerous investigations into the operating charac­ teristics of the engine and its components.

Consequently

10

there exists a considerable quantity of available literature enumerating the results of these investigations.

However,

the review which follows pertains only to the experimentation that lias been carried out on engines built around the u n ­ modified type B turbosupercharger as the compressor-turbine unit. The University of Washington^- has constructed and operated a turbojet engine utilizing the type B-31 turbo­ supercharger as the basic component.

This engine has one

11straight-thro u g h ff can type combustion chamber.

Starting

is accomplished by means of compressed air directed at the exhaust side of the turbine blades through a starting nozzle. The author stated the following as to the purpose of the engine: . . . the use made of our engine will be the investi­ gation and comparison of the operating characteristics of the ’straight-through1 and the 'reverse-flow’ types of combustion chambers under varying primary to second­ ary air ratios and varying flame tube hole sizes and configurations. The effect on efficiency and net thrust with postcompression water injection will also be in­ vestigated. 2 The Northrop Aeronautical Institute^ has constructed

1 Letter of Oliver Poss, University of Washington, William H. Woodward, June 29, 19lf9« 2 Loc. cit. •kt [ _1

3 Report of H. F. Morrison Jr. in the News Views a _j ?__1 *r ..j —_ ~iry |f\17Q I,

to

11

and operated a turbojet engine based on a type B-33 turbo­ supercharger with a view to developing an engine suitable for light aircraft.

This engine includes, in addition to

the turbosupercharger, one tion chamber. employed.

freverse-flow 1 can type combus­

A variable orifice fuel injection nozzle is

The engine is equipped with a variable exit con­

taining a thrust cone.

Starting is accomplished by means

of a high speed electric starting motor coupled to the compressor-turbine

shaft at the intake side of the compressor

by means of a solenoid actuated clutch arrangement.

The

author states that f,the thrust of this school conversion has been cautiously estimated at approximately two hundred pounds. The Cal-Aerotechnical Institute, Glendale, has designed and built a small turbojet engine.

California This engine

is built around an aircraft turbosupercharger of the B-31 type as the basic component.

A single

type combustion chamber is employed.

’straight-flowf can The combustion chamber,

fuel nozzle, and igniter plug are all taken from a Junkers

Jumo-OOij. turbojet engine. thrust cone.

The tail pipe is equipped with a

The engine is brought up to starting r.p.m. by

means of an electric starting motor which engages the turbinecompressor shaft just forward of the compressor inlet. _

12 As yet, performance tests on this engine are incom­ plete.

Theoretical calculations indicate that, with its

present design, this engine should develop a static thrust of approximately 130 lb. with a fuel consumption of approxi­ mately 2.2 lb.

of fuel per minute.

The Marquardt Aircraft Company has produced several turbojet units based on type B turbosuperchargers and is at present making conversion units for use in target planes under Air Force contract. company,

Mr. Edward West, Jr., of this

stated that static thrust of 280 pounds at a specifi

fuel consumption of l.lj-5 pounds per hour per pound of thrust had been obtained at a speed of 26,000 rpm, and a nozzlebox temperature of 1650° F with an air flow rate of 3^2 pounds per minute.

This performance was obtained by redesigning

the turbine nozzles and blades. turbosupercharger,

Without any redesign of the

the maximum thrust of these units was

between lf?0 and l60 pounds.

CHAPTER III DESIGN,

INSTRUMENTATION AND TEST PROCEDURE

As a preliminary step in the analysis of a turbojet engine constructed around the type B-31 turbosupercharger it was necessary to design and construct the complete engine, Essentially the engine consists of the turbosupercharger mounted on a mobile platform with the turbine shaft vertical. Air enters the compressor of the system through an eightsided fairing cone which permits a smooth unobstructed air­ flow.

Pressurized air from the compressor flows through a

stagnation tank designed to divide the air equally between the two combustion chambers.

The gas from the combustion

chamber enters the turbine, expands through the stator blades which direct the exhausting gas against turbine wheel buckets and leaves the system through the vertical exhaust stack. Figure 2 illustrates the completed engine.

Figure 1 is a

schematic diagram showing the numbering of stations. I.

DESIGN AND CONSTRUCTION

Available for use as components on the turbojet engine were the type B-31 turbosupercharger; combustion chamber, fuel booster pump and nozzles from an I-l6 Turbojet engine; electric motors; air compressor; and coils and batteries. fPhft n >*=>« n n m n n n p i r h f l

w a st H^r>. 1 rtp>r! i m n n

14

15

FIGURE

2

OCWLKTK TURBOJET EHGIKE ™

r ° S IT I° "

16

only after preliminary calculations had determined that available components would meet requirements and that fabri­ cation of a new part was not required. Turbosupercharger. No. 75-B-31-A1,

The Turbosupercharger is Model

Type B-31 FW, manufactured by the General

Electric Company in accordance with United States Air Force Specification R-28502-31.

Figure 3 presents a cutaway view

of the turbosupercharger.

Figures ip, 5 and 6 show views of

the turbosupercharger.

This unit consists essentially of a

turbine wheel and centrifugal compressor mounted on the same shaft.

The lubrication system and tachometer leads are con­

tained within the unit. The turbosupercharger and its performance curves com­ prised the basis around which the remaining parts of the turbojet were designed or selected. Intake cone, ducting and exhaust stack.

The intake

cone is enclosed in a large cylindrical tank provided with a screened inlet.

The intake cone entrance was designed to be

approximately six times the cone exit area to obtain low entrance velocity.

Figure 7 is a drawing of the Intake cone

and tank, the ducting and the exhaust stack.

The turbosuper­

charger is included in two views to indicate its relation to the other components.

A stagnation tank, constructed of ten

inch pipe, was used to insure eaual flow to the two combustioi

FIGURE

3

CUTAWAY VIEW OF TURBOSUPERCHARGER

FIGURE

[j.

SIDE VIEW OF TURBOSUPERCHARGER

GENERAL^FLECTR'Cpj

FIGURE

5

FIGURE

6

BOTTOM VIEW OF TURBOSUPERCHARGER

TOP VIEW OF TURBOSUPERCHARGER

SHOWING TURBINE BUCKET WHEEL

SHOWING COMPRESSOR INLET CD

UJNL1LbiJIlU. liLtltL ‘JlY im u riD L

61

au aimsM

J U

Q

k ; *

jlv j.

u x u j *

jh u i i w

u

20

chambers.

Standard five inch pipe was used in ducting from

the stagnation tank to the combustion chamber.

A gradually-

expanding rectangular duct was used from the compressor to the stagnation tank.

The ducts from the combustion chambers

to the turbine, though differing in configuration were designed with the shortest length commensurate with ease of fabrication although this led to higher pressure losses. These pressure losses, less than .3 inches of mercury for the maximum mass flow, although the greatest in any ducting, were too small to warrant the clumsier design that long ducts would have required.

The exhaust stack design was

based primarily on the requirements of safety and measurement of flow conditions.

The gases did not leave the turbine

blades vertically but at an angle which varied with rpm.

In

order to straighten out the flow and obtain uniform condition; in the exhaust stack, vanes and an inner liner 1.8 inches from the stack, sealed at the bottom, were inserted in the stack.

Figure 8 illustrates the exhaust stack. Combustion chambers.

Two combustion chambers from an

I-l6 turbojet were selected to meet the nozzle pressure requirements for good atomization and minimize fuel pump requirements.

The combustion chamber of the I-l6 turbojet

is of the ‘'reverse flow11 type.

Figure 9 is a schematic dia­

gram of the I-l6 combustion chamber.

21

FIGURE

8

EXHAUST STACK

FIGURE

9

QMBUSTION CHAMBER FROM I--l6 TURBOJET

Starting system.

The starting system consisted of an

air compressor; compressed air storage tank; and nozzle, mounted in a sleeve in the exhaust stack, for directing an air jet against the turbine blades.

figure 2 shows the

starting system connected to the engine.

Starts were

obtainable with pressures as low as eighty-five psig. Fuel system.

This system consisted of a fuel pump

and motor, fuel lines and nozzles each of which are discussed below.

Figure 10 is a diagram of the fuel system. The fuel employed during tests on the engine was

eighty octane unleaded gasoline having a higher heating value of 18920 BTU/# and a specific gravity of O.71I4-6 . Of the several available fuel pumps tested the booster pump from the I-l6 turbojet engine most nearly met the re­ quirements of flow rate and pressure.

This hydraulic pump

is rated at 2.98 # gas/min. at two thousand RPM and £0 psi. It is a type DL 19^2 Style A, manufactured by Eclipse Aviation, modified to permit a belt drive. A 220 volt 60 cycle 3 phase, 3/k- HP alternating cur­ rent motor was used to drive the pump at 2600 RPM. Ignition system.

Two Ford "Model T ,f ignition coils

were used to raise the voltage from six volts to l{_000 volts for spark across the spark plugs.

The spark plugs were stan-

y pUSs

Vcj/VC:

t'jii / 't‘(.P y* -- t Vt

I

V

i

.HE i ■ -■ T-ry-

FIGURE

10

SCHEMATIC DRAWING OF FUEL SYSTEM

'Condeno cr~b

V/Cra for C ontucts

Vfbra rin y Co/ /o

FIGURE

11

SCHEMATIC DRAWING OF IGNITION SYSTEM

f\> VA

II.

INSTRUMENTATION

In the analysis of the turbojet engine sufficient instrumentation was required to control the engine, and to record data pertaining to the variables in the gas law and the fuel rate of flow.

This data permitted solution of the

energy equation for conditions existing at each station. This section presents information relative to the instrumen­ tation used. Pressures.

Static and/or stagnation pressure taps

were employed at stations as illustrated in Figure 12. Where more than one pressure tap is located at the same station a single line is drawn from above the manometer board to the station.

A dot is used to indicate manifolding.

It was necessary to accurately measure pressures in each of the five inch ducts leading to the combustion chambers to compare the mass flow to each of the combustion chambers. These values were also compared to the total mass flow through the seven inch intake cone.

In order to measure the

differences between stagnation and static pressures at these stations in inches of water the manometer board was divided into a high pressure system, measuring pressures between the compressor and turbine, and a low pressure system.

A pres­

sure lead from the stagnation tank was manifolded to the high rrnA

cmr»*a

r>o Q P r ’irni r>

a

f1o n o n n o

imn-f-or*

nnl n w n

n ri r)

-hVi ^

Trf crh

27

3*

FIGURE 12 SCHEMATIC DRAWING OF INSTRUMENTATION

pressure side of a mercury U tube.

The other side of the

mercury U tube was open to the atmosphere. the reference levels,

In this manner

an 5*

Conditions were assumed isentropic between stations 0

and 2 and between stations 5 and 6 * measured at stations 2, 3# k-* an 4-* ar*d 5*

The atmospheric temperature and the

temperature used for reference in the Brown Instrument, were measured by ordinary mercury thermometers*

Details of the

instrumentation are discussed in the preceding chapter* The results obtained in the four runs at each turbine speed were extremely uniform, and the conditions of atmos­ pheric temperature, pressure, and fuel rate were the same* Therefore the data were averaged at each rpm*

Based on this

method the recorded data for each turbine speed, together

36 for the analysis, were tabulated in Table I* The computations were made using mechanical computers where possible, and a twenty inch slide rule in all other cases.

They were checked in some instanoes by the use of

five place log tables* Static pressures only were taken at station 3, in the five inch ducts since only the ratio of the flow in the two combustion chambers was desired*

With the two nozzles cali­

brated to pass equal fuel rates the mass flow difference be­ tween the two chambers was less than 1*8^.

The conditions

at station 3 were therefore averaged, and the data consider­ ed as though only one combustion chamber existed* Static and stagnation pressures at station 5(a) were measured by fixed instruments and a survey at quarter inch intervals across the section.

Computations based on the

area average of the survey pressures were compared to values computed by a straight average between the two walls, for stagnation and static pressure*

Since the maximum difference

was less than 0*6%, less than our reading accuracy, the values obtained by straight averages between the wall statics, and stagnation probes one quarter of an inch from each wall were used. The method of computation and the equations used are illustrated by the following example.

Average values obtain-

37 TABLE I TABULATED TEST DATA

Station RPM

Symbol

0-1

2 .2672

3

4

5

6

.250

All

A

10,000

P

w f 1.299

P

Pn 36.75

Ts

12,000

ps

30.090

30.090

1+0 .1+90

39.950 30.211

30.211

w f 1.535

p

30.090

29.896

I+0.300

30.056

30.090

Pn 1*4.5

Ts

14,000

ps

30.090

30.090

l+5 .o!+o

1+1+.271 30.267

30.267

w f 1 .81+1+

p

30.090

29.813

1+1+.796

30.036

30.090

pn 63.75

Ts

16,000

ps

30.090

30.090

5 0 .1+90

1+9 .1+96 30.317

30.317

w f 2 .1I+0

p

30.090

29.726

50.193

30.011

30.090

pn 77.00

Ts

18,000

Ps

30.090

30.090

57.290

56.059 30.336

30.386

Wf 2.1+77

P

30.090

29.602

56.936

29.975

30.090

Pn 91.00

Ts

s

2 (.1363)

30.090

30.090

37 .li|_0

30.090

29.964

36.020

522

561+

522

581

520

603

633

517

515

661+

36.760 30.206

30.206

30.072

30.090

li+20

ii+31

11+33

li+73

1511+

1359

1350

1332

131+0

131+9

38

Station

0

Ps

30.09

P

30.09

t3

1

2

30.09

30.09

3

1*

57.290

5

56.059

30.386

30.386

29.975

30.090

29.602 56.936

515

66J4.

6

1511*.

131*9

Fuel pressure 91 psig Fuel rate

2 •JLj_77 lb/min

Prior to actual computation of results the values of Y

and Cp were determined*

R was assumed constant at

1715 ft lb/slug °R, and values of 2f were taken from “Pro­ pulsion Handbook*“1 * 0 *

y 3= 1 *'+°

cp3 s cPo s *24.Btu/lb °R

y ^ ■ 1 -3^5 cPcc = .2f?3

y 5= 1.355 cPt a *^^4-

Mach numbers were determined from “Propulsion Handbook, M2 = .153

M 3 - .153

- *118

Ratio of temperature across the compressor Ts3/Ts2 = 1.289 Heat increase across the compressor: °Po(Ts3 ' Ts0 ) = 35,8 Btu/lb Heat increase across the combustion chamber: cPcc^Ts4 ~

= 216*5 Btu/lb

1 R. T« DeVault, “Propulsion Handbook,“ Unpublished University of Southern California Report, 20 May 194-9* (Figure 2*) ^ Ibid*

(Figure 3.)

39 Heat loss across turbine: cpt

(Ts^ - T s^) = !j4«l Btu/lb

Ratio of pressure across the compressor:; Pa3/Pg2 = 1.904 Ratio of pressure across the turbine: P 35/^Sk = .542 Pressure loss through combustion chamber as a percentage of Ps3 : 100

— 3-~ .F.g4 3 2.149 % P&3

The adiabatic efficiencies of the turbine and the compressor:

* = s Tr3) Z s f i

-1

w t. /(t

=

= 69*9 *

.749 > 74.9 %

l-(Pa5/Ps4 jV-

The weight flow of air:

w a = 1932

' ,6"

W a = 60 X 32*2

vlb/ min

M = 1932 ( P s / p F 7- a ^ T

at station 2 * W a 3 208.5 lb/min The fuel to air ratio is then W^/Wa = .01187

JjL*M

IfO

The combustion efficiency: v? = cPqc ^Tali_ ~ ts3 ) . .9670 = 96.70 % )(co h (Wf/Wa ) The combined mechanical efficiency of the turbine and com­ pressor system: = °Pc °Pt Tg^)

= .82 =. 82 fo

The thrust: P s mv/ =

W a (l-h W f / W a ) y 1932

M 6"VyRT^ 6 » 22.6 lb (Ts / T ) 6

The specific fuel consumption: c s 60Wf/P s 6#57 lb/hr/lb thrust.

Tabulated results based on similar computations at the other speeds are presented in Table II#

1*1

TABLE II RESULTS OP COMPUTATIONS

RPM Symbols

10,000

12,000

14,000

16,000

18,000

m2

.0776

.0963

.1150

.1320

.1530

M5

.080$

.0862

.1050

.1205

.1530

m6

,01 $0

.0758

.0917

.1034

.1180

106.71

132.1*3

158.60

182.72

208.50

.01217

.01159

.01163

.01171

.01187

Wf/wa

1 .231*.

1 .3l*.6

1.497

1.678

1.904

.8217

.7562

.6837

.6125

.51*20

P/Ps3^

1.020

1 .31*0

1.705

1.969

2.149

T33/Ts0

1.0805

1.1130

1.1600

1.2240

1.2890

Ts5/ T sif

.9570

.91*31*

.9295

.9097

.8910

.775

.781

.765

.712

.699

)?00

.911*5

.9571*

.9385

.9600

.9670

77t

.8775

.7938

.7680

.7653

.7520

"J?E

.595

.660

.743

.790

.820

(HP)t

i*.0.0

66.1

98.7

150.9

211.3

F

9.15

11.04

15.29

20.06

22.60

c

8.53

8.34

7.24

6.39

6.57

•^33/^80

CHAPTER V ANALYSIS OP PERFORMANCE The performance of the engine, as constructed, was below that normally expected of a well designed turbojet* Since no attempt was made to construct an exhaust nozzle the static pressure behind the turbine was essentially atmos­ pheric*

Under these conditions the nozzle box temperatures

were relatively low* ing as a supercharger*

In effect, the engine was still operat­ The analysis of performance was

therefore divided into two sections; first, an analysis of the performance actually obtained, and second, an analysis of the factors causing the low performance as a turbojet* The analysis of actual performance compares the engine to an ideal engine (one w/isentropic compressor — turbine T?cc = 100#) operating under similar conditions*

A

discussion of various operating conditions presented graph­ ically was also included. The analysis of the factors causing the low perform­ ance was presented by component showing how performance could be increased* I*

PRESENT PERFORMANCE

The engine was compared to an ideal engine operat-

1*3

compression ratios were considered the same in each case, permitting solution for conditions behind the compressor. Fuel to air ratio and heat value of fuel were considered the same in each case; stagnation pressure loss through combustion chamber and ducting was neglected# in the nozzle box could then be obtained.

Conditions

With the assump­

tion of one hundred per cent mechanical efficiency the com­ pressor output was set equal to the turbine input#

Condi­

tions downstream of the turbine could then be determined# By expanding the gases from this pressure (assumed to be a stagnation pressure), to atmospheric, it was possible to obtain a theoretical exhaust velocity.

The thrust was then

obtained by multiplying the exhaust velocity by the mass flow of air. The computation of the static thrust produced by the idealized engine, incorporating the above assumptions, was then made as follows: (1 )

Tq3 r T So(TS3/TS o ) s T S o (PS3/PS o )

(2) Energy added to air flow by compressor was f Wa/j?cCP c (Ta3-T3 o ) = W aCPcT So

—i -lj

(3) =T33+ 18920 [(G C

p cc

«T,0 (J» P c C

*=± W f/Wa / P o )? + 18920 P c c

i

+4

(ij.) Energy furnished to turbine to drive compressor

was )?t^P|. (

(5)

)

= W aCp t (Ts^ T 3 5 )

But (2) = (1|_)

therefore £s = i £ps.xhfi T s4 1 CPt H

I"1- (Ps3/PS o ) ^ ] H*-*

x) J

or T 2? = f § ££xT----------- T ^ T ----------- :--- \ ' - W - o > r ] » ] 4 C Pt 3°(rp W Pa,S o )f ^r t I8920w f/w J ^ I8920j £ ^ aa L ^ P cc

(6) (a)

( ^ Y ? 3o

PSo

P s^

PSo

'

PSo

2fc

or (b ) Ps£ _ Ps Psn o " 71

4 ^

Xz±~\ -fo-t ' - l i t CPc/CPt__________________ [ ^ P ^ / P s o ^ J ] /_ , ~gl i Anon ur^,/iiir_ 3 (P a /P s ^ ) ^ 18920 Wf/Wa 'p s 3 Q/-S 3o TSocPco

Prom known operating temperatures it was determined that the following constant values could be used for all engine speeds without appreciable error: GPc s *2^

ftc = l*lf., "Jft* 1 *3 5#

GPcc = *253> Cpt s *26i|_, T So s 520°R

ks Substitution of these values gives

I * - . ! * - . ! * a f i-K—

fi- (£ii )

-

^

V

857

Then static thrust is Wa .Bj^YoRTcr' bo g

(

f -- i--A

1 -Vi-