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The design, construction and instrumentation of a small turbojet engine

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THE DESIGN, CONSTRUCTION, AND INSTRUMENTATION OP A SMALL TURBOJET ENGINE

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 Captain Nelson W. Tobey Lieutenant Almon R. Roth Lieutenant Sylvan E. Salter June 1950

UMI Number: EP60500

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

Dissertation Publishing

UMI EP60500 Published by ProQ uest LLC (2014). Copyright in the Dissertation held by the Author. Microform Edition © ProQ uest LLC. All rights reserved. This work is protected against unauthorized copying under Title 17, United S tates C ode

ProQuest' ProQ uest LLC. 789 E ast Eisenhow er Parkway P.O. Box 1346 Ann Arbor, Ml 4 8 1 0 6 -1 3 4 6

This thesis, written by Captain Nelson W. Tobey Lieutenant Almon R. Roth LeutenanJ. Sylva/i E. f a l l e n der the guidance o/tfe±n.. r acuity Committee, and approved by all its members, has been presented to and accepted by the Council on Graduate Study and Research in partial fulfill­ ment of the requirements fo r the degree of ___

and. guide.*! Missiles.. Date.

June 1950 _______

Faculty Committee

Chairman

TABLE OP CONTENTS CHAPTER I.

PAGE

THE PROBLEM AND DEFINITIONS OF SYMBOLS USED The problem

..

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

Basic principles of a turbojet engine Statement of the problem

1

•• •

3



4

Organization of the remainder ofthe thesis

4

Definitions of symbols u s e d .................

5

REVIEW OF RELATED S T U D I E S ......................

6

Brief history of turbojet power plants • •

••

6



7

Northrop Aeronautical Institute Project 19 University of Washington Project

.........

Marquardt Aircraft Company Project • • • III.

••

8 9

THE MAJOR E L E M E N T S ..............................

11

The B-31 turbo supercharger...................

11

Description

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

11

Specifications ..............................

11

Normal o p e r a t i o n .................. ...

16

The 1-16 combustion c h a m b e r ................. Description and operation IV.

3

••

Importance of the p r o b l e m .............. •

II.

1

.........

Restrictions on the p r o b l e m .........•

1

18

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

18

PRELIMINARY PERFORMANCE ANALYSIS ...............

22

Basis for a n a l y s i s .........................

22

iii CHAPTER

PAGE P r o c e d u r e ..................................

88

Conclusions

33

• . • ..........................

V. THE BASIC DESIGN AND C O N S T R U C T I O N .............

34

Test s t a n d ..................................

34

Compressor i n l e t .................

34

Ducting— compressor to t u r b i n e .............

36

The exhaust s t a c k .........................

38

Assembled e n g i n e ...........................

48

THE AUXILIARY COMPONENTS AND THEIR INSTALLATION

44

VI.

VII.

VIII.

IX.

The fuel s y s t e m ............................

44

The ignition s y s t e m .......................

47

Starting

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

47

L u b r i c a t i o n ................................

49

I N S T R U M E N T A T I O N .................................

56

Measurement of

gas t e m p e r a t u r e s ...........

56

Measurement of

gas p r e s s u r e s ...............

59

Measurement of

r p m .........................

65

ANALYSIS

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

66

Basis for a n a l y s i s .........................

66

P r o c e d u r e .............• • • • ............

66

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

77

DESIGN PERFORMANCE

SUMMARY AND C O N C L U S I O N S .............

80

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

80

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

81

iv CHAPTER

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

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

81 82

LIST OF TABLES TABLE I.

PAGE Turbosupercharger Performance at Various Mass .

Flows and Speeds (Cross Plot D a t a ) II.

III. IV.

Flow Characteristics through the Compressor and Combustion C h a m b e r .. . .......................

32

Corrections for the Shielded Thermocouples • . •

61

Flow Characteristics from Inlet Duct through Compressor .......................

V.

70

........

72

Flow Characteristics from Combustion Chamber Outlet through Turbine . . . . . .

VII.

• • » • • •

Flow Characteristics from Compressor Outlet through Combustion Chamber . . . . .

VI.

29

...........

75

Turbosupercharger Performance at Various Mass Flows and Speeds (Cross Plot Data) . . . . . .

76

LIST OF FIGURES FIGURE

PAGE

1.

Schematic Diagram of a Turbojet Engine

• . . • •

2

2.

Cutaway View of B-31 Turbosupercharger

. . . • •

12 13

Side View of the B-31 Turbo supercharger..... 4m

Turbine End of the B-31 Turbosupercharger • • • •

14

5*

Compressor End of the B-31 Turbosupercharger

15

6.

Schematic Diagram of Turbosupercharged Power

• •

P l a n t .......................

17

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

19

7*

1-16 Combustion Chamber

8.

Components of an 1-16 Fuel Injection Nozzle • • •

9*

Compressor Horsepower Required for a Given Weight

21

Flow and Temperature R i s e .......... 10*

23

Turbine Output Horsepower for a Given Turbine Speed and Pressure R a t i o ..................

24

11*

Compressor D a t a ..............................

25

12#

Correction Factors for Various Fuel/Air Ratios

13*

Cross Plot of Turbine and Compressor Horsepower

*

26

Versus Speed for Various Weight Flows .........

30

14*

Compressor I n l e t ............................

35

15*

Stagnation Tank End P l a t e ...................

37

16*

Ducting— Combustion Chamber to Waste Gate • • • •

39

17*

Dueting--Combustion Chamber to Nozzle Box . . . .

40

18.

Scale Drawing of the Completed Engine ..........

41

vii FIGURE

PAGE

19*

The Turbojet Engine in Testing Position . • . . •

43

20♦

The Fuel S y s t e m ..............................

45

21*

Fuel P u m p ....................................

46

22.

The Ignition S y s t e m .....................

23*

Starting N o z z l e ..............................

24*

Installation of Starting Nozzle .................

51

25.

Schematic Diagram of the Lubrication System . . .

53

26.

Location of Instrumentation.................

57

27.

Components of a Shielded T h e r m o c o u p l e .......



28.

Brown Temperature Recording Instrument

29.

Cross Plot of Turbine and Compressor Horsepower

. . . .

48 50

........

60

Versus Speed for Various Weight Flows......... 30.

Comparative Performance D a t a ...............

78 79

CHAPTER I THE PROBLEM AND DEFINITIONS OF SYMBOLS USED A group of seven graduate students undertook the design, construction, instrumentation, and performance analysis of a small turbojet engine.

The entire group par­

ticipated in all phases of the project, but the recording of the study was divided.

The design, construction, and instru­

mentation are the subject of this report.

The performance

analysis was considered in another report. ^ I.

THE PROBLEM

Basic principles of a turbojet engine.

The arrangement

of the elements of a turbojet engine is shown in Figure 1. The four major elements are the compressor, the combustion chamber, the turbine, and the exhaust nozzle.

The air is

first compressed and then is delivered to the combustion chamber where fuel is added and the fuel-air mixture is ignited.

The hot gases then pass through the gas turbine

which drives the compressor.

The exhaust from the turbine

then discharges through the exhaust nozzle which accelerates the flow.

Provided that the nozzle is ideal, the difference

1 Harold M. Crawford, et al., "The Performance Analysis of a Small Turbojet at Sea-level Conditions,11 (unpublished M a s t e r ^ thesis, The University of Southern California, Los Angeles, 1950).

EXHAUST JET

NOZZLE

FUEL

SCHEMATIC DIAGRAM OF A TURBOJET ENGINE

3 between the momentum at the exit and that at the intake is 2 the thrust produced• Statement of the problem.

The purpose of this project

was to design, construct, and instrument a small turbojet engine using a General Electric B-31 turbosupercharger and one or more 1-16 combustion chambers as the major components* Restrictions on the problem*

The following restric­

tions were imposed: 1.

The engine was to be operated at a speed less than 20,000 rpm and at a nozzlebox temperature below 1400° P* to minimize the probability of a turbine failure*

2.

The engine was to be mounted vertically to eliminate an exhaust jet hazard in a small testing area*

3.

No attempt was to be made to produce thrust; hence no exit nozzle was required.

4*

Maximum use was to be made of surplus and salvage materials available at the University of Southern California.

5*

None of the students assigned to the project were

2 Jet Propulsion (a reference text prepared by the staffs of Guggenneim Aeronautical Laboratory and the Jet Propulsion Laboratory, GALCIT, California Institute of Tech­ nology, for the Air Technical Service Command, 1946), p. 522*

experienced metal workers; hence ease of fabri­ cation was essential. Importance of the problem.

At the present time a

great deal of research is being conducted on gas turbine power plants in general, and in particular on turbojet engines for use in military and commercial aviation.

It is

important, therefore, to have available a small turbojet engine as a teaching aid for engineering students interested in modern engines.

Of equal or greater importance is the

role the engine could play in investigations of various types of combustion chambers, exit nozzles, fuels, and fuel injection devices; and in other problems of importance in the turbojet field. Organization of the remainder of the thesis.

Chapter

II presents a brief history of turbojet power plants and a review of studies related to this project.

The major elements,

the B-31 turbosupercharger and the 1-16 combustion chamber, are discussed in Chapter III, and a preliminary performance analysis based on the characteristics of the major elements is presented in Chapter IV.

Chapters V, VI, and VII describe

the basic design and construction, the auxiliary components and their installation, and the instrumentation respectively. Chapter VIII is a discussion of expected engine performance and Chapter IX is a summary together with conclusions drawn from the study.

5 IX.

DEFINITIONS OF SYMBOLS USED

A - Cross-sectional area, square feet* Cp - Specific heat of air at constant pressure, B.T.U. per pound per degree Fahrenheit* d - Diameter or equivalent diameter, feet* f - Friction factor depending on the Reynolds number and on the type of ducting used, dimensionless• g - Acceleration of gravity (32.2 feet per square second)* H - Heat of combustion of fuel, B*T*U. per pound* L - Length, feet* m - Mass flow, slugs per second, p - Pressure, inches of mercury. Rj) - Reynolds number, dimens ionless* R - Gas constant for air, square feet per square second per degree Fahrenheit.

Value used is 1715.

T - Temperature, degreea Rankine. t - Temperature, degrees Fahrenheit, v -

Velocity, feet per second.

-

Density, slugs per cubic foot.

-

Viscosity of fluid, slugs per foot-second x 10

W -

Weight flow, pounds per minute.



CHAPTER II REVIEW OP RELATED STUDIES The conversion of aircraft superchargers to turbojet engines has been made by several independent groups*

The

most important reason for the conversion is to have available a simple turbojet engine with which studies relative to turbojet engines can be made*

Some designers, however, con­

template its use as a power plant for light planes. Brief history of turbojet power plants.

Early thermal

jet propulsion units had their compressors operated by con­ ventional reciprocating engines.

The concept of a turbojet

was first introduced by Guillaume of France in 1921*

He

suggested that the compressor be driven by a gas turbine, the exhaust of which would constitute the jet.'*' In 1930 Prank Whittle, RAP, filed a patent for a turbojet engine with a multistage axial-radial flow com­ pressor, multiple combustion chambers, a gas turbine, and a o DeLaval nozzle* In 1937 Alf Lysholm received a United States patent

^ Propulsion (a reference text prepared by the staffs of the Guggenheim Aeronautical Laboratory and the Jet Propulsion Laboratory, GALCIT, California Institute of Technology for the Air Technical Service Command, 1946), p. 11. 2 M. J. Zucrow, Principles of Jet Propulsion and Gas Turbines (New York: John Wiley and Sons, Inc., 1948), p. 317.

7 for a turbojet engine with a four stage centrifugal compres­ sor, a single combustion chamber with multiple fuel injectors, a multistage reaction turbine, and a discharge nozzle*

3

The first successful run of a Whittle power plant was made in April, 1937, and the first successful flight was made in England in May, 1941*

The technical data of the

development were turned over to the United States Air Force shortly thereafter*

On October 1, 1942, an American built, 4 turbojet powered plane was flown successfully* Turbojet powered airplanes were used by both the Allies and the Germans in the European Theatre of Operations toward the end of World War II.

The American plane was the

Lockheed XP-80, a reconnaissance plane powered by a General Electric 1-40 turbojet engine.

The Germans used the ME 262,

a twin engine jet fighter. Since the war, development and production of turbojet engines has steadily progressed with applications being made to commercial and military aircraft. Northrop Aeronautical Institute Project 19.

Under the

direction of Faculty Advisor W. L. Tietjen, the students at Northrop Aeronautical Institute have successfully converted

5 Ibid., 'p. 318. 4 "Whittle System of Jet Propulsion," Automotive and Aviation Industries, 90s42, January 15, 1944.

8 a General Electric B-33 turbosupercharger to a small turbojet engine•

The purpose of this conversion was primarily to

familiarize the students with the operation and character­ istics of turbojet engines. The engine was mounted for horizontal exhaust on a test stand equipped to measure static thrust.

It used one

reverse flow type combustion chamber with a jet engine type ignitor, and a variable orifice fuel injection nozzle. Starting was accomplished by a high-speed twenty-four volt direct current electric motor coupled to the compressor by a solenoid actuated dog clutch which could be disengaged after the fuel had been ignited.

The exhaust ducting con­

sisted of an inner tail cone, tail pipe, and interchangeable 5 exhaust nozzles. A similar project using four combustion chambers is now nearing completion. University of Washington Project.

A General Electric

B-31 turbosupercharger has been successfully converted to a turbojet engine at the University of Washington.

This

engine was to be used for the investigation of the operating characteristics of various types of combustion chambers. The unit was started by directing a stream of high velocity air through a nozzle against the turbine blades.

The air

5 H. P. Morrison, Jr., flTurbo Conversion Design,” News Views (Los Angeless Northrop Aeronautical Institute, December 15, 1948)•

was furnished from a thirty-five cubic foot tank, which, with a pressure of 160 pounds per square inch, accelerated the turbine to 5000 rpm with a pressure drop of only 20 pounds per square inch.

The space between the compressor

and turbine was enclosed with a sheet metal strip through which low pressure air (five to ten pounds per square inch) was circulated to cool the unit#

6

Marquardt Aircraft Company Project.

Mr. Edward West,

Jr., of the Marqyardt Aircraft Company, Van Nuys, California, has produced several turbojet engines from General Electric type B turbosuperchargers.

At present these conversion units

are being produced under Air Force contract for use in target planes.

Preliminary designs have been made by Mr. West for

a personal plane powered by two turbojet engines converted from turbosuperchargers. Mr. West stated that by redesigning the turbine nozzles and blades, it has been possible to attain a static thrust of 280 pounds at a specific fuel consumption of 1.45 pounds per pound of thrust per hour, a speed of 26,000 revo­ lutions per minute, a nozzlebox temperature of 1650° F., and an air flow rate of 342 pounds per minute.

With a redesign

of the compressor, he anticipates a thrust as high as 400

6 Oliver Foss, faculty, University of Washington, personal letter to Captain W. H. Woodward, July, 1949.

10 pounds.

It was his opinion that, without any redesign of

the turbosupercharger unit, thrusts in the neighborhood of 150 to 160 pounds should be attainable.

CHAPTER III THE MAJOR ELEMENTS The General Electric B-31 turbosupercharger and the 1-16 combustion chambers were considered the major com­ ponents of the small turbojet engine.

An explanation of the

normal method of operation of each was considered adequate, since their use in the conversion is essentially that for which they were designed. I.

THE B-31 TURBOSUPERCHARGER

Description.

Essentially, the B-31 turbosupercharger

is a variable speed, centrifugal type air compressor driven directly by a gas turbine, which is in turn normally driven by the energy of the exhaust gas from an aircraft engine. Figure 2 is a cutaway view of the B-31 turbosupercharger. The impeller of the compressor appearing in the upper part of the photograph is shafted directly to the turbine wheel below.

The lubrication pump, geared off the main shaft,

appears at right center.

Other views of the unit are shown

in Figures 3, 4, and 5. Specifications.

The normal rated characteristics of

the B-31 turbosupercharger are 5^ 1 Supercharger Installation Manual, GET-1002A (Schenec tady, New York: General Electric Company, 1944), Sec. B, p. 3

FIGURE 2 CUTAWAY VIEW OF B-31 TURBOSUPERCHARGER

H

to

FIGURE 3 SIDE VIEW OF THE B-31 TURBOSUPERCHARGER

H

14

FIGURE 4 TURBINE END OF THE B-31 TURBOSUPERCHARGER

15

G E N ERAL

ELECTRIC ^

FIGURE 5 COMPRESSOR END OF THE B-31 TURBOSUPERCHARGER

16 Approximate weight (incl. accessories), lb.

144 24,000

Rated speed, RPM

1600

Nozzlebox temperature, °F.

120

Weight flow, lb./min. Compressor discharge pressure, in. Hg. abs*

31.67

Compressor inlet pressure, In. Hg. Abs. NACA altitude, ft. Normal operation.

10 28,000

A turbosupercharger enables an

aircraft engine to perform efficiently at high altitudes by compressing the thin atmosphere of the upper regions to approximately sea-level density for delivery to the carburetor, p thereby maintaining normal intake manifold pressure* Figure 6 is a schematic diagram of a turbosupercharged power plant illustrating normal design operation of the B-31 turbosupercharger.

The hot exhaust gases from the engine are

ducted to the nozzlebox of the turbosupercharger, through the turbine wheel or out the waste gate.

The power thus

imparted to the turbine is transmitted by a shaft to the compressor.

The rotational speed of the turbine, and hence

the power delivered to the compressor, is controlled by a device that automatically positions the waste gate so that 2 Operation, Service, and Overhaul Instructions for Turbosupercharger Types B-2, B-ll, B-22, B-51, and B-33 (Chicago: Marshall-White Press, 1945), p. 20. 3 Ibid., p. 21.

Exhaust

INTERCOOLER

BOOST CONTROL

RAM MiNG

manifold

AIR < iNTakf -

CX tank

CARBURETOR

ELECTRIC REGUl atqr

' INTERNAL supercharger

-

INTAKE

MANIFOLD

ex h a u st

sta ck

-

NOZZLEBOX



EXHAUST GASES

[

10 I L

H H

jwASTEGATE^

COMPRESSED AIR



_

L OVERSPEED + W GENERATOR

__

ATMOSPHERIC AIR

FIGURE 6 SCHEMATIC DIAGRAM OF TURBOSUPERCHARGED POWER PLANT

18 the carburetor intake pressure remains constant for the desired engine performance at all altitudes.

Atmospheric

air is supplied to the compressor through a ramming air in­ take and is delivered to the carburetor at essentially sealevel pressure.^ II.

THE

Description and chamber which

1-16 COMBUSTION CHAMBER operation.

The 1-16 eombustion

was used on the turbojet is a reverse-flow,

annular type,combustion chamber shown

diagrammatically in

Figure 7. The air from the compressor enters the combustion chamber from the left and flows in the annular space between the outer casing and the inner flame tube.

Part of the air

continues through the annular space to the end of the combustion chamber where it makes a 180 degree turn and mixes with the fuel injected by the 1-16 fuel nozzle.

The

rich mixture created is then ignited by the spark from the ignitor plug.

The very hot products of combustion move to

the left inside the flame tube and are cooled by the remainder of the cold air which is introduced through the holes in the flame tube.

This final mixture then makes a 90 degree turn

to enter the turbine nozzlebox.

The many turns in flow

direction make the mixing complete but also cause the 4 Ibid., p. 22.

FIGURE 7

1-16 cohbustioi: ci:a”BSR

pressure drop to be very high.

5

The fuel injection nozzle used with the 1-16 com­ bustion chamber is standard.

Its components are shown in

Figure 8#

5 Jet Propulsion (a reference text prepared by the staffs of the Guggenheim Aeronautical Laboratory and the Jet Propulsion Laboratory, GALCIT, California Institute of Technology for the Air Technical Service Command, 1946), p. 441.

3JJOT

2par»

B03T

^

V

?tcs

FIGURE 8 COMPONENTS OF AN 1-16 FUEL INJECTION NOZZLE

10

H

CHAPTER IV PRELIMINARY PERFORMANCE ANALYSIS A preliminary analysis based on ideal operating con­ ditions was conducted to determine,

(1) an operating point

for each of several arbitrarily assumed air flow rates and an assumed turbine nozzlebox inlet temperature, and (2) the design required to insure sufficient fuel flow for operation at the point corresponding to the maximum air flow rate* Basis for analysis*

The graphs of Figures 9, 10, and

11 were taken from charts furnished by the General Electric Company

and are for sea-level conditions only*

The General

Electric Company charts were our only source of data on the turbosupercharger performance and, since they were designed for application to conventional aircraft power plant operation (Cf* Chapter III, p. 16), they were not sufficient for a complete analysis of the turbojet problem (Cf* Chapter I, pp* 1-3)*

The charts were based on a nozzlebox inlet temper­

ature of 1400° F* whereas the actual turbine nozzlebox inlet temperature could be expected to vary through a wide range, depending on the fuel rate.

Further, the maximum air flow

rate shown on the charts was 200 pounds per minute, whereas 1 Supercharger Installation Manual, GET-1022A (Schenectady, New York: General Electric Company, 1944), Sec, 0, p. 6, Figure 0-4*

23

ilL80

m

;:r

o 1A0

SJlO'O

£

i 20

10 60 80 100 120 1X0 160 Compressor Horsepower Required

180

FIGURE 9 C0TTRH°S0R HORSEPOWER REQUIREL FOR A GIVEN WEIGHT FLOW AND TEMPERATURE RICH

200

o

2 u 6 8 10 12 IK 16 10 Equivalent Turbine ap««d - Thousands RRI FIGURE 10 TURBINE OUTPUT HORSEPOUER FOR A GIVEN TURBINE SPEED AND PRESSURE RATIO

20

trgwj s^ead (at 37 T lnltft tdnpersta Oto£ aosjogs comisressor iure

2.5

2.0

OQ0^

io,ooq

1.0 20

0

40 60 80 100 120 130 140 Souivalant Vfelght Flow at 59 F Ihltt Tamperature • #/min 1*)00 2000 Inlet Volume Flew at 59 F Inlet Temperature - cu, Ft./rain.

FIGURE 11 COMPRESSOR DATA

200

220

3000

Correction

Factor

26

Fuel/Air Ratio FIGURE 12 CORRECTION FACT0R3 FOR VARIOUS FULL/AIR- RATIOS

27 there was a possibility that the maximum rate for the turbojet would be considerably above 200 pounds per minute (Cf* Chapter II, p* 9). The basic assumptions made were: 1*

Approximately standard sea-level conditions at the inlet to the compressor:

pressure, 30 inches

of mercury; temperature, 60° F.j density, 0*002378 slugs per cubic foot* 2*

A nozzlebox inlet temperature of 1400° F. to agree with the General Electric Company charts*

3.

Heat of combustion of the fuel, 18,920 B.T.U. per pound.

4*

Specific heat of air at combustion chamber eonditions, 0*27 B.T.U* per pound per degree Fahrenheit.

5*

Ideal combustion chamber efficiency*

6.

No pressure losses in ducting or combustion chamber.

7*

No heat losses due to radiation.

8*

Uniform conditions across each cross-sectional area*

9. 10.

Atmospheric pressure at the turbine outlet. A certain air flow rate for each point of operation*

2 George A. Hawkins, Thermodynamics (Third printing; New York: John Wiley and Sons^ Inc•, 1947), p* 95*

p

23 Procedure.

The initial problem was to select a rotor

speed for each assumed mass flow such that the turbine horse­ power was just sufficient to operate the compressor.

This

was accomplished by a trial and error method resulting in the cross plot of turbine horsepower and compressor horsepower versus rotor speed (see Figure 13).

These cross plots were

made for assumed air flow rates of 100, 120, 140, 160, 180, and 200 pounds per minute. For each assumed air flow rate and trial rotor speed selected, Figure 11 yielded the pressure ratio and temperature rise factor across the compressor.

From these two quantities

the pressure and temperature at the compressor outlet were determined.

Using the temperature rise through the compressor

and the assumed mass flow, the horsepower required to operate the compressor was read from Figure 9. The turbine performance charts were constructed for a fuel/air ratio of 0.075 based on the aircraft engine operation. However, the literature indicated much lower fuel/air ratios for turbojet operation.

Taking 0.02 as an average fuel/air

ratio, an approximate correction factor of 1.02 was obtained 4 from Figure 12. Multiplying this factor by the trial rotor 3 Jet Propulsion (a reference text prepared by the staffs of Guggenheim Aeronautical Laboratory and the Jet Propulsion Laboratory, GALCIT, California Institute of Tech­ nology, for the Air Technical Service Command, 1946), p. 528. 4 Supercharger Installation Manual, op. cit., Sec. 0, p. 9, Figure 0-6.

TURBOSUPERCHARGER PERFORMANCE AT VARIOUS MASS FLOWS AND SPEEDS (Cross Plot Data)

Mass Flow #/min

Pressure Ratio Across Turbine

Equivalent Turbine Speed, RPM

Compressor Horsepower Required

Turbine Output HP

47

1.26

10,200

27

20

.140

73

1.41

12,240

41

74

10,000

•085

44

1.24

10,200

32

14

12,000 10,000

.135 .085

70 44

1.40 1.20

12,240 10,200

48 34

72 4

12,000

.135

70

1.38

12,240

57

67

12,000

.130

68

1.32

12,240

64

53

14,000 14,000

.165 .165

86 86

1.48 1.41

14,280 14,280

76 90

100 85

16,000 16,000

.220 .220

114 114

1.64 1.58

16,320 16,320

117 130

144 128

18,000

.290

151

1.91

18,360

173

216

Comp* Speed RPM

Temp* Rise Factor Across Compressor

10,000

*090

12,000

Temp. Rise Across Comp, op

100

120

140

160

180

200 Notes

Multiply compressor speed by 1*02 to get equivalent turbine speed.

20

t-1t■

19 — I_

Speed

- Thousands

RPM

IS 17

i:

16 15 14 13

11

10 40

60

80

;:;i±H r* rU} 100 120

140

160

180

200

220

240

260

280

Horsepower FIGURE 13 CROSS PLOT OF TURBINE AND COMPRESSOR K0RSEP0..ER VERSUS SPEED FOR VARIOUS Y/EIGHT FLOPS

VjJ

o

31 speed gave an equivalent turbine speed* With the assumptions of atmospheric pressure at the compressor inlet and turbine outlet, and no pressure losses in the ducting or combustion chamber, the pressure ratio across the turbine is equal to the reciprocal of the pressure ratio across the compressor.

For the equivalent turbine

speed and the pressure ratio across the turbine, the horse­ power output of the turbine was read from Figure 10* The compressor and turbine horsepower were determined in this manner for several trial rotor speeds at each assumed air flow rate*

These values are plotted in Figure 13*

The

intersection of the two curves at any given air flow rate was designated as an operating point* The compressor outlet temperature was determined for each operating point, using the temperature rise factor corresponding to the rotor speed at the operating point. The temperature rise across the combustion chamber was found by subtracting the compressor outlet temperature from the assumed nozzlebox temperature* The weight flow of fuel required to produce the indi­ cated temperature rise across the combustion chamber was computed using the equation

(4.1) where A T chamber.

Wfuel * W&

is the temperature rise across the combustion

TABLE II FLOW CHARACTERISTICS THROUGH THE COMPRESSOR AND COMBUSTION CHAMBER

Weight Flow Temp, at Operating Temp. Rise Temp. Rise Temp, at Turbine Weight Flow Factor Across Across Comp. Speed Nozzlebox of Fuel Into of Air at Comp. Temp. Comp. Inlet Inlet RPM Compressor Outlet Compressor Comb. Chamber Ojp °p °F #/min. #/min. 100

60

10,620

.106

55

115

1400

1.832

120

60

11,200

.114

59

119

1400

2.195

140

60

11,860

.127

66

126

1400

2.545

160

60

13,000

.146

76

136

1400

2.885

180

60

14,800

.183

95

155

1400

3.200

200

60

16,750

.243

126

186

1400

3.465

33 Results ©f these calculations for the operating points at various air flow rates are shown in Table II. Conclusions.

The results of the calculations indi­

cated that two 1-16 combustion chambers would be necessary, since the fuel flow required to operate at the higher air flow rates and rotor speeds was greater than the maximum capacity of one 1-16 combustion chamber fuel nozzle.

CHAPTER V THE BASIC DESIGN AND CONSTRUCTION The principal factors that dictated the design of the small turbojet engine from a General Electric B-31 turbosupercharger were (1) the design requirements of a maximum speed of 20,000 rpm and a maximum nozzlebox temperature of 1400° F.;

(2) the availability of 1-16 burners for the com­

bustion chamber; and (3) the restricted area assigned for operation.

The preliminary analysis indicated that, to meet

the design requirements, two 1-16 burners would be necessary to pass the required fuel flow.

The exhaust hazard in the

small area available indicated that a vertical exhaust would be necessary.

The basic design problem was then to design a

small turbojet engine using a General Electric B-31 turbo­ supercharger, two 1-16 burners, and a vertical exhaust. Test stand.

A mobile test stand which was available

from the Mechanical Engineering Department was adapted to fit the mounting bolts of the turbosupercharger.

Sand boxes

were mounted on the test stand to minimize the effects of flying fragments in the event of a turbine failure. Compressor inlet.

The compressor inlet (Figure 14)

was designed to minimize the possibility of drawing foreign materials into the compressor, and to provide a seven inch

’EES n 9I/T JO JHOO ’ES!G W K E 9I/I JO SBTH m h o n 91s k i h b i H i : HI I HVDS

m a hossmm n

a

mm s m j

36 straight section in which to straighten the flow sufficiently for accurate static pressure measurements to be used in the computation of mass flow* Due ting*--compress or to turbine.

The use of two

burners presented the problem of passing equal mass flow of air from the compressor through each burner*

A stagnation

tank, made from a 55*5 inch section of ten inch steel pipe, was introduced between the compressor and the burners to solve the problem.

End plates equipped with a thermocouple

and a pressure tap (Figure 15) were bolted to each end of the stagnation tank. The ducting from the compressor to the stagnation tank was mad© from one eighth inch steel and installed as shown in Figure 18. To minimize the ducting from the burners to the nozzle­ box and the waste gate, it was necessary to join the inlet side of the burners to the stagnation tank with a 20.5 inch and a 20.0 inch (centerline measurements) section of five inch inner diameter steel pipe for the waste gate side and nozzlebox respectively.

The stagnation tank ends of these

pipes were cut to fit the contour of the stagnation tank and welded in place.

Static pressure taps were installed at the

mid-length sections.

The combustion chamber inlet adapter

plates were welded in place to complete the ducting from the stagnation tank to the combustion chambers.

37

FIGURE 15 STAGNATION TANK END PLATE

38 The ducting from the combustion chamber to the waste gate was made from one sixteenth inch stainless steel*

The

ends of the ducting were'welded to the combustion chamber outlet adapter plate and the waste gate adapter plate, and the completed assembly was bolted into position as shown in Figure 16*

Three static pressure taps, one movable total

pressure tube, and three thermocouples were Installed* The ducting from the combustion chamber to the nozzle­ box is shown in Figure 17*

The instrumentation and materials

used for this duct were the same as those used for the ducting to the waste gate side of the turbine*

The ducting was

welded to the combustion chamber outlet adapter plate and the nozzlebox adapter plate and bolted in position as shown in Figure 18* The exhaust stack.

Since no attempt was to be made

to produce thrust with the small turbojet engine, a straight section of pipe was originally used for an exhaust stack*

A

rectangular piece of one eighth inch steel 42 by 43 inches was rolled Into a cylinder and welded along the seam*

Holes

were drilled in the stack to accept the Instrumentation and the starting nozzle sleeve*

The stack was then welded to the

turbine exhaust adapter plate and bolted Into position* This did not prove to be a satisfactory design, since the flow was much too turbulent to obtain suitable stagnation

39

FIGURE 16 DUCTING--COMBUSTION CHAMBER TO WASTE GATE

FIGURE 17 DUCTING--COMBUSTION CHAMBER TO NOZZLE BOX

o

u

♦--X—

SCALE

1 IN - 18 IN

FICUHE IS SCALE DRAWING OF THE COMPLETED ENGINE

42 pressure measurements•

To overcome this difficulty, a cylin­

drical central body with vertical vanes was inserted in the original exhaust stack.

This annular type of exhaust stack

and vane arrangement straightened the flow sufficiently to yield usable pressure measurements. Assembled engine.

Figure 18 is & I t l Q

(one to

eighteen) scale drawing of the completed engine, and Figure 19 shows the engine in testing position.

43

& Jt

i i 0

FIGURE 19 the

turbojet

engine

in

testing

position

CHAPTER VI THE AUXILIARY COMPONENTS AND THEIR INSTALLATION The fuel, ignition, starting, and lubrication systems were designated as the auxiliary components•

Their operation

and installation are discussed below* The fuel system. matically in Figure 20.

The fuel system is shown diagramThe fuel tank was a five

with a petcock and hose connection at the bottom.

gallon drum The unit

was mounted on a scale to determine the fuel flow during each test run.

Refueling of the tank was accomplished by gravity

feed from a twenty-five gallon reservoir. Figure 21 is a photograph of the fuel pump and the fuel pump motor.

The former was an Eclipse Aviation hydraulic

pump rated at one half gallon per per minute and 50 pounds

minute at 2,000 revolutions

per square inch pressure.

The fuel

pump motor used was a three quarter horsepower, 220 volt, three phase, electric motor with a V-belt drive.

The combi­

nation was tested and found to be capable of delivering the required fuel flow.

The fuel pump motor switch was located

on the control panel. The fuel flow was controlled by the by-pass needle valve during normal operation.

An emergency cut-off valve

was placed in the fuel line as indicated and mounted on the

,j/ oas^ Vo/ve

/ in

FIGURE 22 TEE IGNITION SYSTEM

49 Two possible starting devices were a high-speed electric motor and a high-velocity air jet (Cf. Chapter II). The air jet was used, employing a sixteen cubic foot air storage tank certified at 125 psia.

For a nozzle, an opening

one inch by one eighth inch was formed on the end of a one inch diameter steel pipe as shown in Figure 23.

The nozzle

was installed on the low pressure side of the turbine wheel at an angle of about thirty degrees to the plane of rotation as shown in Figure 24, and was connected to the release valve of the air storage tank by a one inch inner diameter Vulco rubber hose (Figure 19). With this starting system the turbosupercharger could be accelerated to 5000 rpm in about six seconds.

During this

time the storage tank suffered a loss of about 20 psia.

A

speed between 4000 and 5000 rpm could be maintained for about twenty-five seconds with a loss of about 50 psi.

This proved

to be satisfactory for starting. Lubrication.

Since a turbosupercharger operates at

very high rotational speeds (and at high altitudes in aircraft), the lubrication problem must be given careful consideration to insure correct operation. The lubrication system consisted essentially of a supply tank, supply and return tubing, supply pump, and scavenger pump.

The turbosupercharger had its own built-in

supply pump and scavenger pump as a compact, high-speed,

50

FIGURE

23

STARTING NOZZLE

FIGURE 24 INSTALLATION OF STARTING NOZZLE

Cn H

, 52 double-gear-type unit.

This unit, called the lubricating

pump, was really two separate, positive displacement pumps on the same shaft.

It was geared to the main shaft of the

turbosupercharger.

The lubricating pump shaft was geared

down from the main shaft by a ratio of ls9.5. The bearing-and-pump casing as shown schematically in Figure 25 was located between the nozzlebox and the com­ pressor casing, being securely bolted to both.

Lubricating

oil from the supply pump (1) entered the inside of the bearingand-pump casing through a shroud.

The oil was delivered by

jets (3) to the mesh of the drive gear and the worm thread sleeve (4) and to the ball and roller bearings (5).

The

bearings were also oiled by the mist which existed inside the bearing casing as a result of the high rotational speeds and churning of the oil.

The combination of splashing and oil

mist was ideal lubrication for the ball and roller bearings. The jets which delivered oil directly on the ball and roller bearings provided no better lubrication than the oil mist, but did provide more efficient cooling of the bearings and l permitted higher operating speeds. The oil which collected In the bearing-and-pump casing was removed through a sump bushing by the larger gear pump, called the scavenging pump (1), the capacity of which was 1 Supercharger Installation Manual (Schenectady, New Yorks General Electric Company, 1944), Sec. H, p. 1.

VA typcs

e-n. e si — ► B 22 B 8 53

FIGURE 25 SCHEMATIC DIAGRAM OF THE LUBRICATION SYSTEM

cn 01

54 about three times that of the supply pump.

Because of this,

two thirds of the scavenging pump delivery was air.

The

pumping of this air caused a slight vacuum in the bearing housing, which was necessary to prevent oil leakage through the shaft oil seals.

The two shaft oil seals, one in the

turbine end and the other in the compressor end of the bearing housing, were not rubbing seals, but had a clearance from the shaft.

These seals were threaded to cause an inward

flow, which tended to keep the oil inside the pump-and-bearing casing.

This action was assisted by the vaeuum which was

present inside the casing as described above.

The dumbell

valve (6) operated by gravity and was to provide for proper lubrication regardless of the position of the turbosupercharger in flight.

2

Since the supply pump and scavenger pump were built-in, the only parts of the lubrication system which had to be added were the supply tank and supply and return tubing. Minimum specifications for this part of the system as recom­ mended by General Electric Company are listed below and were 3 adhered to in the design of the turbojet engine*

2 Ibid ." 7 Sec. H, p. 2. 3 Ibid., Sec. H, p. 3*

55 Oil Tank Size, Gal. 2.0

"Pull Tank" Rated Plow Tube Gal, ______ GPM_____ Diam. 1.5

1

Oil-in

(P.P.)

Grade of Oil

3/4 in.

5580-D

Oil-out 1/2 in.

The excess volume of the tank indicated by the differ­ ence between oil tank size and "full tank" above was necessary to accommodate any foaming of the oil induced by the turbosupercharger scavenging pump.

CHAPTER VII INSTRUMENTATION The instrumentation consisted of measuring the speed of the turbosupercharger rotor and of measuring the gas temperatures and pressures at critical locations throughout the engine• Measurement of gas temperatures .

A total of nineteen

thermocouple pyrometers were used for temperature measurement in the engine (Figure 26), three in each of the two entrance ducts to the nozzlebox (Figures 16 and 17), five in the upper exhaust stack, six in the lower exhaust stack, and two in the stagnation tank (Figure 15).

Seventeen of these

thermocouples were where the temperatures would be high enough to require shielding for accuracy.

Shielding was not

used on the two stagnation tank thermocouples. Common mercury thermometers were used to measure the atmospheric temperature at the engine intake and in the Brown instrument at the cold junction of the thermocouples. The National Bureau of Standards has made an extensive study to develop thermocouple pyrometers for gas turbines and has recommended several reliable types applicable to this project.'*'

The type shown disassembled in Figure 27 was used

1 Eighth Monthly Report of Progress on the Development of Thermocouple Pyrometers for Gas Turbines~Twashington, D.C.i National Bureau of Standards, 1946), 15 pp.

57

T*

FIGUIiE 26 location of

m sm m

58

Porcelain Insulator

No. 22

A luce 1 Wire

Junction for

Fused Junction

O.D. Stainless Steel Tube

O.D. Stainless Steel Tube

FIGURE 27

COMPONENTS OF A SHIELDED THERMOCOUPLE

leads

59 in the engine because it is accurate, easy and inexpensive to fabricate, and easy to install*

Corrections shown in

Table III were reported by the National Bureau of Standards to be accurate within 4° P. for all values up to 20° P. and to within one fifth of the correction for all higher values.

o

These corrections do not apply, of course, to the stagnation tank unshielded thermocouples* Insulated chrome1-alumel leads were taken from the head of each thermocouple to a male coupling (on© to each side of the test stand) where a flexible steel cable carried chrome1-alumel leads on to the terminals of a Brown continu­ ous balance recorder (Figure 28).

Millivolt readings from

the Brown instrument were corrected, depending on the local cold junction temperature, to an equivalent reading for a 32° F. cold junction temperature before being converted to temperatures in degrees Parenheit.

3

Measurement of gas pressures»

The performance analysis

of the small turbojet engine necessitated the measurement of stagnation and static pressure on the upstream and downstream side of each major component of the engine.

These measure­

ments were used in conjunction with temperature measurements

2 Ibid., Table III* 3 Installation, Operation, and Maintenance, Electronic Instruments (Philadelphia; Brown Instrument Company, Sec. 1137, pp. 9-10.

60

v • • mmm «■ • •MB

FIGURE 28 BROWN TEMPERATURE RECORDING INSTRUMENT

61

TABLE III CORRECTIONS FOR THE SHIELDED THERMOCOUPLES

Thermoeoupl© Reading

Wall Temp.

Corrections to Thermocouple Readings for the Following Gas Flows in Lb/sec ft^ 2 4 6 8 Degrees Fahrenheit

1500 1500 1500 1500 1500

1500 1400 1300 1200 1100

0 13 24 33 41

0 8 15 20 25

0 6 11 15 18

0 5 9 12 15

1400 1400 1400 1400

1400 1300 1200 1100

0 11 20 28

0 7 12 17

0 5 9 12

0 4 7 10

1300 1300 1300 1300

1300 1200 1100 1000

0 9 17 23

0 6 10 14

0 4 7 10

0 3 6 8

1200 1200 1200

1200 1100 1000

0 3 14

0 5 9

0 3 6

0 3 5

1100 1100 1100

1100 1000 900

0 6 11

0 4 7

0 3 5

0 2 4

1000 1000 1000

1000 900 800

0 5 9

0 3 6

0 2 4

0 2 3

62 at corresponding points to compute the mass flow, the adiabatic efficiencies of the compressor and turbine, the compression ratio of the compressor, the combustion efficiency of the burners, and the theoretical thrust of the engine* The measurement of static pressure in a flowing fluid can be made with adequate accuracy at a solid wall along which the fluid flows, by inserting a small metal tube flush with the inside wall at the point where the measurement is to be made, and connecting the other end of the tube to a manometer board.^ This system was used for the static pressure taps throughout the engine* Figure 26*

These were located as indicated in

In each case one eighth inch stainless steel

tubing was inserted through the ducting wall, silver soldered in position, and smoothed with a file and emery eloth to remove any burrs at the mouth of the tube*

The existence of

burrs and other imperfections could have led to completely false pressure readings.

5

The use of wall pressure taps on the exhaust stack and inner body as a means of measuring static pressure in

the

turbulent flow of that region requires justification.

Before

4 I*. Prandtl and 0* G* Tietjens, Applied Hydro-and Aero-Mechanics (New Yorks McGraw-Hill Book Company, Inc*, 1934), p. 227* 5 Ibid., p. 228*

installing these taps, a survey of static pressure was made using a static tube as a probe.

Readings were taken, over

the operating range of r p m fs, at one quarter inch intervals across the exhaust.

In each position the probe was rotated

until the minimum pressure was registered and this pressure was then recorded.

It was decided that a wall tap on the

inner body and outer body would be sufficient to record the exhaust static pressure.

This was later confirmed on actual

test runs by taking readings from both the probe and wall 7 taps • The stagnation pressure of the entering air was taken to be atmospheric and read from a mercury barometer.

An

atmospheric reference level was maintained as indicated in Figure 26 on both the low and high pressure sides of the manometer board.

The pressure reading from the stagnation

tank end plate tap (Figure 15) was used as the stagnation pressure between the compressor and the burners. In the ducting between the burners and the turbine inlets, stagnation pressure probes were Installed.

These

were made from one eighth inch stainless steel tubing ten inches long, counter bored one eighth inch in from the probe 6 Ibid.. p. 227. 7 Harold M. Crawford, et al., f,The Performance Analysis of a Small Turbojet at Sea-level Conditions,11 (unpublished Master's thesis, The University of Southern California, Los Angeles, 1950).

64 end, and bent at right angles one inch from the same end# Each was inserted in the ducting, through a sleeve equipped with a lock nut (Figure 16) to secure it in any desired position#

The position of each was adjusted to give a

maximum pressure reading for each operating rpm. A stagnation pressure survey was made across the annular exhaust area at quarter inch intervals with a stag­ nation pressure probe similar to the one described above# Maximum pressure was obtained at each position by rotating the probe.

The survey indicated that four stagnation pressure

tubes similar to those installed in the ducting to the turbine would be required, and that these should be installed in sets of two on opposite sides of the exhaust stack, inclined at an angle of fifteen degrees from the vertical.

In each set, one

tube was placed one quarter of an inch from the inner wall, and the other, one quarter inch from the outer wall, and silver soldered in position. To make use of the rubber tubing available, a short piece of one quarter inch copper tubing was silver soldered over the outside end of each pressure tube#

These were then

connected by rubber tubes to a bank of twenty-five two inch sections of copper tubing numbered consecutively for easy identification and attached to the engine test stand#

The

other end of the copper tubes were connected through rubber tubing to the manometer board.

Measurement of rpm.

The General Electric B-31 turbo-

supercharger used in the conversion was ©quipped with an exterior tachometer cable fitting driven by the oil pump* (See (4), Figure 25)*

The ratio of tachometer speed to

turbine wheel speed was It9*S.

Since the control panel was

to be some distance from the engine during operation, a standard aircraft tachometer generator and dial combination were used*

The generator was mounted on the engine and con­

nected to the tachometer drive by a flexible tachometer cable. The dial was mounted on the control panel and electrically connected to the generator.

CHAPTER VIII DESIGN PERFORMANCE ANALYSIS The preliminary analysis of the turbojet discounted entirely all pressure losses in the ducting and combustion chamber, and assumed maximum combustion efficiency.

It is

the purpose of this chapter to determine operating points for the same assumed air weight flows considering the pressure losses and combustion chamber efficiency. Basis for analysis♦

In this analysis the General

Electric Company charts were used in the same manner as in Chapter IV. The basic assumptions made were the same as in Chapter IV excepts

(1) pressure losses occurred in the

ducting and combustion chambers, and (2) non-ideal combustion chamber efficiency prevailed. Procedure.

The following station designations were

used throughout the analysis: Station 0 - scoop inlet Station 1 - compressor inlet Station 2 - compressor outlet Station 3 - combustion chamber inlet Station 4 - combustion chamber outlet Station 5 - turbine inlet

67 Station 6 - turbine outlet Velocities for weight flows of 100, 120, 140, 160, 180, and 200 pounds of air per minute were computed at station "0" by the equation (7*1)

v * m./jOk

The area at station W0 W was 1.668 square feet and the density was taken as 0.002378 slugs per cubic foot. The Mach number (M) at station ”0 ” was determined by the equation (7.2)

M =* v/a

where "a" is the local speed of sound.

Since the Mach number

and the area were known at station M0 M, the local sonic area could be determined and was assumed to be constant through the duct.

1

Knowing the sonic area and the area at station

W1M, the Mach number at station Ml ” was determined.

Equation

7.2 then gave the velocity at station ”1 ”. Pressure losses between stations 0 and 1 were caused by friction and the decrease in cross-sectional area.

The

former losses were dependent on the type of cross section, roughness of the duet, length of the duct, number and types of bends, the degree of contraction or expansion of the duct, and the temperature, density, and velocity of the flow.

The

^ C. L. Dailey and F. C. Wood, Computation Curves for Compressible Fluid Problems (New York! John Wiley and Sons, Inc., 1949), Figure 1-5.

68 equation used for computing the pressure loss due to friction was

2 (7*3)

AP * i ^ f L .

Since the contraction of the inlet scoop was gradual, and the air flow was at very low subsonic speeds, it was assumed that the section was a straight pipe having a mean diameter of 0.987 feet and a length of 0.75 feet. The friction factor, ftf tf, was obtained from tables which were entered with the type of material of the ducting and the Reynolds number. 3

The type of material used was

welded steel and the Reynolds number was computed using the 4 equation (7.4)

R^ * />vdf/K .

The pressure loss in the section of straight duct between the inlet scoop and the compressor was computed in the same manner.

This duct was 0.841 feet long and 0.583

feet in diameter. The pressure loss due to friction was computed, as outlined above, for both ducts for each air weight flow, 2 John R. Weske, Pressure Losses in Ducts with Compound Elbows (Washington, D.C.s National Advisory Committee for Aeronautics, ARR, February, 1943). 3 Glenn Murphy, Mechanics of Fluids (Scranton, Pennsylvanias International Textbook Company, 1943), p. 140. 4 Ibid., p. 136.

69 combined, and entered in Table IV as the -

P ducting,

station ff0 ,f - ”1 ”. The pressure loss due to the decreasing crosssectional area was determined by equation (7*5) '

2

where P s is the stagnation pressure.

The values for Pg/P at

stations "0" and ”1" were determined from a plot of the above g

equation.

Since the pressure at station ”0 ” was known, and

since P s was assumed to be constant, the pressure loss was computed for each weight flow of air and entered in Table IV as -

P contraction, station ,f0 ” - Mlff. The

subtracted

combined pressure losses through the ducting from the original

were

pressure and entered under the

column Station r,l rf in Table IV along with the velocity at station "1” . The

pressure ratio and temperature rise factor across

thecompressor were

obtained

as explained in Chapter IV and

entered in Table IV, in the column Station Hl ” - ”2 ”.

These

were used to compute the pressure and temperature at station ”2 ” for each weight flow and rotor speed. The density at station "2” was computed from the equation of state 5 Dailey, 0£. cjLt., p. 4. 6 Ibid., Figure 1-1.

TABLE IV FLOW CHARACTERISTICS FROM INLET DUCT THROUGH COMPRESSOR

Station n^»t

Station «0m _

Station w0"

Station «1« _ «t2it

RFM W

P

t

V

- A? Ducting

- A? Contraction

P

.10

29.394

V

Comp. Ratio

Temp. Rise Factor

1.26

.09

1.41

.14

1.24

.085

1.40

.135

1.38

.135

14,000

1.50

.171

12,000

1.32

.13

14,000

1.48

.165

14.000

1.41

.165

16.000

1.64

.22

16,000

1.58

.22

1.91

.29

10,000 100

30

60

13*92

.006

76

12,000 10,000 120

30

60

16*70

.009

.15

29.841

95

12,000 12,000 140

160

180

200 18,000

30

30

30

30

60

60

60

60

19.43

22.25

25.05

27.80

.011

.015

.018

.023

.20

.30

.35

.45

29.789

29.685

29.632

29.527

115

134

142

162 o

The velocity at station f,2,f, where the area was 0.272 square feet, was computed using equation 7.1, Tor the various weight flows of air being considered* The ducting from the compressor outlet to the stag­ nation tank was a short section of semi-rectangular ducting 0.835 feet long with an equivalent diameter of 0.579 feet* The pressure loss due to friction in this section was com­ puted as above using equations 7.3 and 7.4 assuming zero velocity in the stagnation tank. The ducting from the stagnation tank to the combustion chamber inlets consisted of two ducts 0.417 feet in diameter. One was 1.71 feet long, and the other 1.67 feet long. Pressure losses were assumed equal for each and computed for the longer pipe using equations 7.3 and 7.4.

The velocities

used were computed on the basis of equal mass flows in each duct using equation 7.1, with the density computed at station 2, and entered in Table V. The sum of the two ducting losses computed above was entered in Table V as the - A p —

Ducting, station ”2” - f,3M*

Conditions at station W3 W were entered as the result of the above computations assuming that the temperature remained constant, and the mass flow was equally divided between the two combustion chambers.

TABLE V FLOW CHARACTERISTICS FROM COMPRESSOR OUTLET THROUGH COMBUSTION CHAMBER

Station W2W - tt3w

Station ”2 tt

Station ”5”

Station "3lt-,t4l

RPM w

P

t

V

P— Ducting

W

P

t

Y

-4 P Comb. Chamber

57.65 107

69.2

.017

50

57.655

107

69.5

1.47

42.10 155

64.5

.016

50

42.084

153

65.0

1.64

56.95 104

85.7

.022

60

56.926

104

84.6

1.44

12,000

41.70 150

77.4

.025

60

41.677

150

78.3

1.63

12,000

41.00 150

92.2

.051

70

40.969

130

92.7

1.60

14,000

44.60 149

101.5

.055

70

44.567

149

102.0

1.78

12,000

59.10 128

110.2

.042

80

59.058

128

110.8

1.52

146

101.7

.059

80

45.761

146

102.1

1.75

41.70 146

120.0

.049

90

41.651

146

120.8

1.67

16,000

48.50 174

107.9

.045

90

48.455

174

108.7

2.22

16,000

46.60 174

124.6

.057

100

46.545

174

125.0

2.14

56.40 211

109.6

•050

100

56.550

211

109.7

2.59

10,000 100 12,000 10,000 120

140

160 o CO • to

14,000 14,000 180

200 18,000

73 Pressure losses across the combustion chamber were taken from available empirical data and entered in the table as - A P comb, chamber.

7

Conditions at station ”4 ” were entered in the table assuming that the temperature was 1400° P., and neglecting the amount of fuel added in the combustion process.

Equation

7.1 was used to compute the velocity, since the area at station ”4 ” was known to be 0.0875 square feet, and the density could be computed by equation 7.6. In making the computations for station ,f5 w , it was considered sufficiently accurate to compute only the changes across the ducting leading to the regular nozzlebox entrance, rather than to attempt to arrive at an average value based on the losses in the two similar duets. The area at station ”5 ” being 0.21 square feet, the velocity was computed by the same methods used to compute the velocity at station nl ”. The pressure change between stations "4" and w5 ff due to the expanding duct, was computed by the same methods used to compute the pressure change for the contracting section between stations ”0 ” and "1” . there was a pressure increase.

However, in this instance, The values found were entered

7 Ray E. Boltz and John B. Meigs, Fuel Tests on an 1-16 Jet Propulsion Engine at Static Sea Level Conditions (Washington, D.C.t National Advisory Committee for Aero­ nautics Reference Memorandum No. E7B01), Figure 15.

74 in Table VI as

^ p — expansion*

Friction losses were not

taken into account, since they are negligible whenever the 0 expansion angle is greater than fifteen degrees* The conditions at station ”5 ” were inserted in the table for the various rotor speeds and air weight flows, and a pressure of thirty inches of mercury inserted in the column station W6M in each case to complete the table* The temperature rise across the compressor was computed for each case from Table IV and entered in Table VII* The pressure ratio across from the pressure at station ”5 ”

the turbine was computed and station W6W in Table VI

and entered In Table VII* The combustion efficiency of the 1-16 combustion 9 chamber was found to be approximately 97 per cent# Hence, the fuel/air ratio could be computed by the relationship (7.8)

% u e l „ F/A „ c w air *

a t /.97 h .

£ T was computed in each case as the difference in the temperature between stations w3 n Table VII*

and ”4 ” and entered in

It was then used In equation 7*8 to compute the

next column of the table, the fuel/air ratio.

An approximate

correction factor of 1*02 was read from Figure 12 and 8 Supercharger Installation Manual, GET-1022A (Schenec­ tady, New York: The General Electric Company, 1944), Sec. E, p. 5* 9 Boltz,

ojd.

cit., Figure 14.

TABLE VI PLOW CHARACTERISTICS FROM COMBUSTION CHAMBER OUTLET THROUGH TURBINE

Station H4tt

Station Wglf

Station ff4n - w5n

Station "6"

RPM w

P 36.163

10,000

40.444

12,000

35.486

10,000

40.047 39.369

12,000

42.787 37.538

12,000

42.001 39.991

14.000

46.235 44.429

16,000

.637

36.80

331

.456

40.90

450

1.014

36.50

53.760

V

400

.853

40.90

474

1.231

40.60

30 134 181 30 165 190

1400 434

1,213

44.00

569

1.762

39.30

30 175 227

1400 534

1.599

43.65

601

2,109

42.1

30 214 239

1400 517

1.865

48.1

600

2.271

46.7

30 206 237

1400 497

1.840

55.6

P

148

1400

1400

100

t 1400

1400

90 16.000

367

1400

80 14,000

P

1400

70 14,000

+ A P Expansion

1400

60 12,000

V

1400

60

18,000

t

30 198

TABLE VII TURBOSUPERCHARGER PERFORMANCE AT VARIOUS MASS FLOWS AND SPEEDS (CROSS PLOT DATA)

Comp. Weight AT Pressure AT Across Fuel/ Speed Equiv. Comp. Turbine Speed Flow Across Ratio Across Comb. Air Correction Turbine Horsepower Horsepower RPM of Air Comp. Turbine Chamber Ratio Factor Speed Required Output ______ #/min._____________________________________________________________________________ 47

1.23

1293

.0190

1.02

10,200

27

15

12,000

75

1.36

1267

.0186

n

12,240

41

65

10,000

44

1.22

1296

.0190

n

10,200

32

14

12,000

70

1.36

1270

.0187

tt

12,240

48

64

12,000

70

1.35

1270

.0187

tt

12,240

57

60

14,000

89

1.47

1251

.0184

tt

14,280

72

96

12,000

68

1.31

1272

.0187

tt

12,240

64

47

14,000

86

1.46

1254

.0184

tt

14,280

76

94

14,000

86

1.41

1254

.0184

tt

14,280

90

78

16,000

114

1.61

1226

.0180

tt

16,320

117

127

16,000

114

1.56

1226

.0180

tt

16,320

130

120

151

1.86

1189

.0175

tt

18,360

173

200

10,000 100

120

140

160

180

200 18,000

77 multiplied by the compressor speed to give an equivalent turbine speed. The horsepowers required to operate the compressor and delivered by the turbine were evaluated, plotted, and an operating point determined as in Chapter IV (Figure 29)* Conclusion.

The plots of the design performance

analysis and preliminary performance analysis operating points (Figure 30) indicated that the effect of the pressure losses was small (approximately 5 per eent reduction of air weight flow at maximum rpm) ♦ suitable.

Hence, the ducting design was

KPM - Thousands Speed 0

20

40

60

80

100

120

140

160

Horsepower FIGURE 29 CROSS PLOT OF TURBINE AHI COMPRESSOR HORSEPO'.'EE VERSUS SPEED FOR VARIOUS HEIGHT FL04S

^

100

120

1^0

160

Air Weight Flow - ^/rain, FIGURE 30 C0TTPARAT IVE PERFORMANCE DATA

180

200

CHAPTER IX SUMMARY AND CONCLUSIONS The design, construction, and instrumentation of a small turbojet engine using a General Electric B-31 turbosupercharger and two 1-16 combustion chambers as the major components was successfully completed* Summary*

The principle factors that dictated the

design of the small turbojet engine were (1) the design requirements of a maximum speed of 20,000 rpm and a maximum nozzle box temperature of 1400° F.,

(2) the availability of

1-16 combustion chambers, and (3) the restricted area assign­ ed for operation*

A preliminary analysis Indicated that, to

meet the design requirements, two 1-16 combustion chambers would be necessary to pass the required fuel flow*

The

exhaust hazard In the small area available indicated that the engine should b© mounted to exhaust vertically* An expected performance analysis conducted after the design phase showed that the efficiency of the 1-16 combustion chambers and the pressure losses through the ducting and com­ bustion chambers should not greatly alter the performance characteristics of the engine.

Therefore, the engine was

installed as designed. The completed engine operated smoothly during approxi­ mately twelve hours of intermittent testing, and was operated

continuously for about two hours on several occasions.

The

instrumentation proved to be satisfactory for a quantitative performance analysis. Conclusions, (1) Two 1-16 combustion chambers were necessary to pass the required fuel flows at the higher rpm's, (2) The pressure losses through the combustion chambers and ducting should cause only about a 5 per cent reduction in the air weight flow, (3) A small turbojet engine can be built, operated, and performance tested using the design and instrumentation presented in this thesis. Recommendations,

The following recommendations are

made: (1) That the small turbojet engine be used as a teach­ ing aid in jet propulsion courses, (2) That an exit nozzle be installed and thrust calculations be made, (3) That the engine be used in investigations of various types of combustion chambers, ducting, exit nozzles, fuels, and fuel injection devices; and in other problems of importance in the turbojet field.

BIBLIOGRAPHY

BIBLIOGRAPHY Boltz, Ray E., and John B. Meigs, Fuel Tests on an 1-16 Jet Propulsion Engine at Static Sea Level Conditions* Washington: N.A.C.A., Ref* Memo* No* E7B01* Crawford, Harold M*, et al., "The Performance Analysis of a Small Turbojet at Sea-level Conditions." Unpublished Master*s thesis, The University of Southern California, Los Angeles, 1950. Dailey, C* L., and F. C. Wood, Computation Curves for Com­ pressible Fluid Problems* New York: John Wiley and Sons Inc., 1949* 33 pp. Eighth Monthly Report of Progress on the Development of Thermocouple Pyrometers for Gas Turbines, Washington, D.C. s National Bureau of Standards, 1946* 15 pp. Foss, Oliver, personal letter to Captain W. H* Woodward, Seattle, Washington, July, 1949* Fries, Stuart Gilbert, John Blair Beach, and Wilbur Leonard Kahn, "The Operation and Limitations of the 1-16 Com­ bustion Chamber*11 Unpublished Master's thesis, The University of Southern California, Los Angeles, 1949* 46 pp. Hawkins, George A., Thermodynamics. New Yorks John Wiley and Sons, Inc., Third Printing, 1947. 436 pp. Installation* Operation* and Maintenance of Electronic Instruments* Philadelphia: Brown Instrument Company* Jet Propulsion* A reference text prepared by the staffs of the Guggenheim Aeronautical Laboratory and the Jet Propulsion Laboratory, GALCIT, California Institute of Technology for the Air Technical Service Command, 1946. 799 pp. Morrison, H. F*, 'Jr., "Turbo Conversion Design,11 News Views* Los Angeles: Northrop Aeronautical Institute, December, 1948. 12 pp. Murphy, Glenn, Mechanics of Fluids* Scranton, Pennsylvania: International Textbook Company, 1943. 329 pp.

84 Operation, Service, and Overhaul Instructions for Turbo­ supercharger Types B-2, B-Tl, B-22, B-51, and B-33. Chicago: Marshall-White Press, 1945, 218 pp. Prandtl, L , , and 0. G. Tietjens, Applied Hydro- and AeroMechanics. New Yorks McGraw-Hill Book Company, Inc., 1934. 311 pp. Supe rcharger Installation Manual GET-1002 A . Schenectady, New Yorks General Electric Company, 1944. 200 pp. Weske, John R., Pressure Losses in Ducts with Compound Elbows. Washington, D.C.S National Advisory Committee for Aeronautics, ARR, February, 1943. "Whittle System of Jet Propulsion," Automotive and Aviation Industries, 90:42, January 15, 1944. Zucrow, M. J., Principles of Jet Propulsion and Gas Turbines. New Yorks John Wiley and Sons, Inc., 1948. 563 pp.

U n ^ i n l t j o f S o u t h e r n C a lif o r n ia L lbfW V