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DOCUMENT NO.__

This is copy number J 'of~, which consists of

44 . A__ pages, series _ •

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April 1959

(TITLE UNCLASSIFIED)

By

T. R. Ward and James NeelyETF, ARO, Inc. 2;I'Pdi f ' r; .J ',' /' r~';

. +.f; : ;''':- i'

ARNOLD ENGINEERINGDEVELOPMENT CENTER

AUTHORITY: AEDC/PA 2001-151

Approved For Public Release, Distribution Is Unlimited.(Declassified Dec 1959 per Bill Moss June 19,2001).

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ALTITUDE PERFORMANCE EVALUATION OF TWOAFTERBURNER FUEL MANIFOLD CONFIGURATIONS

IN AN IROQUOIS SERIES 2 TURBOJET ENGINE

DECLASSIFIED - UNCLASSIFIED /,J'

AEDC- TN-59-11ASTIA DOCUMENT NO.:AD·306443

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ARMED SERVICES TECHNICAL INFORMATION AGENCYARLINGTON HAll STATIONARLINGTON 12, VIRGINIA

ATTN: TISVV

Department of Defense contractors must be

established for ASTIA services, or have

their need-to-know certified by the cogni­

zant military agency of their project or

contract.

~l.UXbt

(Title Unclassified)

AEDC·TN·59·11ASTIA DOCUMENT NO.:

AD·306443

ALTITUDE PERFORMANCE EV ALUATION OF TWO

AFTERBURNER FUEL MANIFOLD CONFIGURATIONS

IN AN IROQUOIS SERIES 2 TURBOJET ENGINE

By

T. R. Ward and James Neely

ETF# ARO# Inc.

April 1959

System - Canadian CF 105 (Arrow Interceptor)

Project 3066 - Task 30532ARO Project No. 121841

Contract No. AF 40(600)-700 SI A 13(59-1)

AEDC- TN-59-11

. . . .

CONTENTS

ABSTRACT •••NOMENCLATUREINTRODUCTION .APPARATUSPROCEDURERESULTS AND DISCUSSION •SUMMARY OF RESULTSREFERENCES . • . • • •APPENDIXES

A. Methods of Calculation • • • . • • •B. Tabulated Data (limited distribution under separate

cover). . . . . • . . . • • . • • . • . • • . . • .

ILLUSTRATIONS

Figure

1. Iroquois AX-102 Engine Prior to Installation in TestChamber . . . . • • • . . • . • . •

2. Internal View of Iroquois Afterburner

Page

4477

10121415

16

25

26

3. Afterburner Fuel Manifolds . . . . • . 27

4. Installation of Iroquois in Altitude Test Chamber 28

5. Test Article Station Locations ....••.. 29

6. Instrumentation Location Details (looking upstream) . 30

7. Iroquois Series 2 Operational Envelope. . 32

8. Afterburner Test Data Obtained . • •. .••. 33

9. High Pressure Compressor Surge Limit . • • • 34

10. Afterburner Fuel Manifold Flow Characteristics ..•. 35

11. Afterburner Stability Limits. • • . • • • • • . . • .. 36

12. Afterburne~ Combustion Efficiency. . 39

13. Engine Pumping Characteristics ..••••.•. 40

14. Generalized Compressor Airflow . • •. .•. 41

15. Non-afterburning Engine Performance at Mach Number1. 47 at Altitude . . • . • • • • • • .' • • • • • • •• 42

AEDC-TN-S9-11

. . . .

Figure

16. Afterburning Engine Performance at Mach Number1. 47 at Altitude . . . . • • • • .• .•..

17. Exhaust Nozzle Discharge Coefficient

Page

43

44

3

AEDC- TN-59-11

ABSTRACT

At the request of Headquarters, United States Air Force, an investi­gation was conducted to determine the relative performance of two after­burner fuel manifold configurations in an Orenda Engines Ltd. Iroquoisseries 2 turbojet engine. Performance data were recorded at a constantengine inlet temperature of 5600 R and at a constant high pressure rotorspeed of 7650 rpm for engine inlet pressures ranging from approximately2 to 8 psia. The results indicate that the performance of the single entrymanifold configuration was equal to or better than that of the triple entrymanifold configuration at most of the conditions tested.

Limited compressor performance data were obtained at a Reynoldsnumber index of O. 37 with exhaust nozzle areas ranging from 4.5 to 5.28sq ft. Some overall engine performance results are also included.

NOMENCLATURE

A

a

CF.Jis

Area, ft 2

Speed of sound at simulated flight altitude

Exhaust nozzle discharge coefficient

Flow coefficient

Ratio of actual jet thrust to jet thrust produced by anideal convergent nozzle expansion of gas at jet pipeconditions to test chamber condition

Ratio of actual jet thrust to jet thrust produced byisentropic expansion of gas at jet pipe conditions totest chamber conditions

Thrust, lb

Jet thrust obtained from scale force, lb

Scale force

Fuel-air ratio

Simulated flight inlet momentum" lb

Acceleration due to gravity" 32.14 ft/sec2

High pressure compressor inlet guide vane position, deg

Enthalpy, Btu lIb

Lower heating value of fuel, Btu lIb

4

M

m

P

P

PLP

R

ReI

RF

SFC

T

t

V

W

y

SUBSCRIPTS

a

ab

adj

AEDC-TN-59-11

Mach number

Mass rate of flow, lb-sec /ft

Rotational speed of low pressure compressor-turbine rotor,rpm

Rotational speed of high pressure compressor-turbine rotor,rpm

Total pressure, psfa, psia

Static pressure, psfa, psia

Power lever position, deg

Gas constant, 53.34 ft-lb/lb-oR

Reynolds number index

Thermocouple impact -recovery factor

Specific fuel consumption

Total temperature, oR

Static temperature, oR

Velocity, ft/sec

Weight rate of flow, lb/sec, lb/hr

Ratio of specific heat at constant pressure to specificheat at constant volume.

Ratio of absolute total pressure to absolute totalpressure at sea-level static conditions of ARDC modelatmosphere (2116 psfa)

Change in enthalpy, Btu /lb

Efficiency

Ratio of absolute total temperature to absolute totaltemperature at sea-level static conditions of ARDCmodel atmosphere (518. 7°R) .

Ratio of absolute viscosity to absolute viscosity atsea-level static conditions of ARDC model atmosphere

Air

Afterburner

Related to simulated flight conditions

5

b

bl

c

cc

cn

e

eff

f

g

hp

i

is

j

lp

m

n

s

t

tp

tu

x

00,0,12,3,etc

1n

lOs

00

AE DC- TN-59-11

Burner

Station 4 overboard bleed

Compressor

Combustion chamber

Ideal convergent exhaust nozzle

Engine

Effective

Fuel

Gas

High pressure compressor

Indicated

Isentropic

Jet

Low pressure compressor

Pressure loss due to heat addition

Net

Scale

Total

Tailpipe

Turbine

Effective duct outside diameter in labyrinth seal plane

Instrumentation station locations

Venturi throat

Exhaust nozzle static pressure

Equivalent free - stream conditions

6

AEDC-TN-59-11

INTRODUCTION

At the request of Headquarters" United States Air Force" a testingprogram was conducted at the Engine Test Facility" Arnold EngineeringDevelopment Center (ETF-AEDC)" to evaluate the altitude performanceof the Iroquois turbojet engine manufactured by Orenda Engines Ltd.Malton" Ontario" Canada.

The two objectives of the portion of the test reported here wereto determine compressor performance and to determine the relativemerits of two afterburner fuel manifold configurations used on theIroquois engine. All testing was accomplished on a single Iroquois .turbojet engine (SiN AX-102) during the period September 5 throughOctober 13, 1958.

Since the second objective was the more important, it is treatedmore thoroughly in this report" and the results were more thoroughlydocumented. Only limited data were obtained pertaining to the firstobjective" and only typical performance plots are presented. Some over­all engine performance results are also included. A summary of themechanical operating experiences accumulated during the testing ispresented in Ref. 1.

APPARATUS

ENGINE

The Orenda Iroquois series 2 engine (Fig. 1 and Ref. 2) is a t.wo­spool axial flow turbojet engine with a close coupled afterburner de­signed for supersonic flight. It has a mechanical limitation of Machnumber 2. 0 at 38" 000 -ft altitude at hot day conditions or Mach number2.13 at 36" OOO-ft altitude at standard day conditions.

The low pressure compressor-turbine spool consists of a threestage compressor driven by a single turbine stage. The high pressurecompressor-turbine is made up of a seven stage compressor driven bya two stage turbine. Estimated sea-level static military thrust of theengine without afterburning is 18, 570 lb at a low pressure rotor speedof 5740 rpm and a high pressure rotor speed of 7930 rpm (~ef. 3).

Manuscript released by authors February 1959.

7

AEDC- TN-59-11

The engine combustion system is composed of an annular combustionchamber and a vaporizing fuel system with 32 equally spaced fuel injectionpipes.

The Iroquois afterburner (Fig. 2) consists basically of a flame stabi­lizing system, a short perforated internal liner immediately downstreamof the gutters, two fuel manifolds, a jet pipe with an external shroud onthe downstream portion only, and a fully modulating petal-type convergingexhaust nozzle. The flame stabilizing system has an inner and an outerring gutter with a fuel manifold immediately upstream of each gutter. Asecond afterburner fuel manifold is located approximately 13 in. upstreamof the downstream fuel manifolds and radially midway between them.

The two afterburner fuel manifold configurations tested were thesingle entry fuel manifold configuration and the triple entry fuel mani­fold configuration. In addition, some data were taken during operationwith the downstream manifold only.

The single entry fuel manifold configuration consisted of a singleentry upstream fuel manifold (Fig. 3a) and the downstream fuel mani­fold (Fig. 3c). The triple entry fuel manifold consisted of a triple entryheat-shielded upstream manifold (Fig. 3b) used with the down$treammanifold (Fig. 3c).

The single entry upstream manifold (Fig. 3a) is an oval cross­sectional tube having orifices drilled in the upstream edge for fuelspraying and only a single fuel entry line located at approximately the1800 position. The triple entry upstream fuel manifold (Fig. 3b) isheat-shielded, has three fuel entry lines, and utilizes a round cross­sectional tube with orifices facing upstream for fuel injection. Thedownstream fuel manifold (Fig. 3c) consists of ten separately fed inde­pendent manifolds of round cross-sectional tubing with orifices locatedin both the downstream edge and upstream edge for fuel inj ection. Theupstream and downstream manifolds were supplied from the samesource and operated at the same pressures when used simultaneously.

Engine and afterburner fuel flows were manually controlled.

The position of the exhaust nozzle during the engine checkout andengine evaluation portions of testing was manually controlled. Theposition of the exhaust nozzle during afterburning was controlled auto­matically to maintain a constant ratio of compressor discharge pres­sure to turbine discharge pressure of 3. 2.

The engine was equipped with inlet guide vanes upstream of thehigh pressure compressor rotor; their position was manually controlledto meet the needs of the test.

8

AEDC- TN-S9-11

INSTALLATION

The engine and air measuring venturi were installed in the altitudetest chamber (Fig. 4 and Ref. 4), and the engine was attached to theexit of the air measuring venturi by me ans of a flexible boot. The frontof the venturi, which was supported from the thrust stand with theengine, was suspended inside the labyrinth seals, which were attachedto the rear face of the plenum chamber bulkhead, thus giving a "free­flo ating" unit.

Air was supplied through the cooling air manifold to maintain anindicated test chamber air temperature less than 610 0 R.

INSTRUMENTATION

Aerodynamic pressures and temperatures were measured at thestations shown in Fig. 5 with instrumentation as shown in Fig. 6.

The multiple probe rakes at stations 2, 3, 4, 7, 9, and 9a weredesigned to measure conditions at the centers of equal areas. Pres­sures and temperatures were measured at station 9a until the station9a rake was removed for the afterburner testing. Inlet plenum pres­sure was indicated from two static orifices. Test chamber ambientpressure was indicated by four static orifices in the test chamber.

Aerodynamic pressures were recorded by photographing manometerboards in which mercury, tetrobromoethane, and water were used asindicating fluids. Temperatures were indicated on millivolt recordersand converted to degrees by a digital computer. The temperatureswere tabulated in the control room immediately after each data run,and some specific performance parameters were tabulated to permitimmediate evaluation of the validity of the data run.

Scale force measurements were obtained by the flexure-pivot-typethrust stand reacting against a strain-gage load cell mounted In the testchamber. Rotor speeds were indicated on events-per-unit-time counters.Liquid flows were sensed by turbine-type flowmeters and indicated onevents-per-unit-time counters. Flowmeter calibrations were madeusing JP-4 fuel.

The nozzle position was sensed by a potentiometer and indicated bya null balance potentiometer system.

Observation of the burning characteristics during afterburningoperation was accomplished through a periscope mounted in the testchamber exhaust ducting. In addition, color motion pictures were madeof the burning characteristics; a copy is on file in the ETF-AEDC.

9

AE DC- TN·59·11

Certain compressor blade stresses, engine vibrations, the exhaustnozzle area, and engine rotor speeds were recorded on oscillographsduring periods when engine speed, afterburner fuel flow, and engineinlet conditions were being changed.

PROCEDURE

Air at the temperature and pressure desired was supplied to theengine inlet. Performance was related to the simulated flight condi­tions which were defined by the engine inlet conditions, Aircraft Indus­tries Association (AlA) inlet recovery, and the ARDC model atmosphereby means of the choked nozzle technique (Ref. 5). The test chamberambient pressure surrounding the engine was set as near the desiredaltitude ambient pressure as the plant capabilities would allow.

In those cases in which the test chamber ambient pressure couldnot be set at the proper level to simulate the desired altitude, the over­all performance was related to the desired altitude by an adjustmentto the thrust values if the exhaust nozzle was choked. This adjustmentis the product of the difference in pressure and the effective nozzle areaand may be expressed as the ratio of effective velocities obtained withthe desired exhaust nozzle pressure ratio to that obtained with the actualexhaust nozzle pressure ratio (See Appendix A). The parameters usingthis thrust are noted with the subscript "adj".

Certain parameters have been standardized to sea-Ievel-standardconditions by the use of 0 and () (as defined in the Nomenclature) and arereferred to as generalized.

The theoretical overall performance of the engine equipped with ahypothetical convergent-divergent nozzle with a throat area equal tothe exhaust nozzle area and an exit area equal to 12. 4375 sq ft was tobe calculated if the tailpipe pressure were high enough to operate ondesign or underexpanded at the altitude conditions simulated. Sincethis was never the case, the calculation was not accomplished.

The testing was accomplished in three parts - engine checkout,engine performance evaluation, and afterburner evaluation.

ENGINE CHECKOUT

The engine checkout portion of testing was performed to determinethat the engine and installation were functioning properly. Operationat 8 psia inlet total pressure and 5600 R inlet total temperature was

10

AEDC-TN-59-11

possible, but the engine could not be operated at 4 psia engine inlettotal pressure because of compressor surge.

COMPRESSOR PERFORMANCE EVALUATION

The compressor performance portion of the testing was limited toonly two high pressure compressor rotor speeds (6000 and 6500 rpm) ateach of four set exhaust nozzle areas between 4. 65 and 5. 28 sq ft. Opera­tion at higher high pressure rotor speeds was not possible because ofcompressor surge. During this portion of the test, engine inlet totalpressure was 4. 0 psia, and engine inlet total temperature was 4150 R.

Data were also obtained to indicate the engine surge limit in theoperating area where the afterburner investigation was to be made usingzero, 1.17 percent, and 2.34 percent high pressure compressor dis­charge overboard bleed. (Normal Iroquois engine operation utilizes ahigh pressure compressor discharge bleed of approximately 2 percentto drive the fuel pumps. )

AFTERBURNER EVALUATION

The afterburner evaluation portion of the testing was accomplishedat the simulated flight conditions shown in Fig. 7. The Iroquois series 2operating envelope is based on data presented in the model specification(Ref. 3). For all afterburning testing, engine inlet total temperaturewas 560 0 R, high pressure compressor rotor speed was 7650 rpm, andhigh pressure compressor inlet guide vane position was 0 deg.

Data were first obtained by using the single entry upstream fuelmanifold configuration; random data runs were also made with the up­stream manifold out of service, during which time operation was withthe downstream manifold only. The single entry upstream fuel mani­fold was then replaced with the triple entry upstream manifold, andsimilar data runs were made. At each inlet pressure, data wereobtained between lean and rich blowout or the limiting nozzle area. Asummary of the afterburner data runs is shown in Fig. 8.

Data obtained to. determine afterburner stability and blowout limitsconsisted of measurements of afterburner fuel flow and turbine dischargetotal pressure. These values were recorded manually from control roomobservations, and a full steady-state data point was not taken.

Engine power is identified by high pressure compressor rotor speed.The engine was operated within the manufacturer's tolerances for AEDCoperation of this engine (Ref. 6).

11

AE DC-TN-59-11

Fuel used throughout the testing conformed to Mil-F-5624C gradeJP-4 with a lower heating value of approximately 18" 600 Btu/lb. Oilused throughout the testing conformed to Mil-L-7808C.

The methods of calculation used in the reduction of the data aregiven in Appendix A.

The data obtained during the test are tabulated in Appendix B(under separate cover). An explanation of the identification code isgiven on the first page of the Appendix. Identification of the columnheadings is given in this report.

RESULTS AND DISCUSSION

COMPRESSOR PERFORMANCE EVALUATION

The results obtained from the compressor performance evaluationportion of the testing are decidedly inconclusive in respect to the exactlevel of performance over a wide range of engine rotor speeds andexhaust nozzle areas. At each exhaust nozzle area set" an engine surgeoccurred before a high pressure rotor speed of 7000 rpm could bereached.

The limitation thus placed on the range of operation was so re­trictive that first one" and then two" overboard bleed ports 1. 92 in. indiameter were opened to bleed high pressure compressor discharge air.Bleed flow rate through each port was approximately 1. 17 percent ofcompressor airflow. An indication of the effectiveness of this over­board bleed is shown in Fig. 9. The high pressure rotor speed at whichstall would occur with the 2. 34-percent overboard bleed was 6900 rpmat the minimum area tested" which is a generalized rotor speed of 7714rpm. Since the highest generalized high pressure rotor speed to beattained during afterburner evaluation was 7363 rpm (7650 rpm at 5600 RLthis margin was felt to be safe.

AFTERBURNER EVALUATION

The afterburner fuel manifold flow characteristics are shown inFig. 10. Ideally the relationship of flow rate to pressure drop of anincompressible fluid through an orifice should have a slope of O. 5 ona logarithmic plot. This was the case for all manifold configurationsat the high flow rate end of the curves; however" as the flow rate wasreduced" first the single entry configuration and then the triple entryconfigurations began to deviate. The downstream manifold did not

12

AE DC-TN-59-11

display any deviation from the straight line above a flow rate of approxi­mately 4000 lb/hr and was quite close down to approximately 3000 lb/hr.

Since the gage used to indicate the upstream manifold pressure wasconnected to the fuel line which fed the manifold, rather than the mani­fold itself, there was some change in indicated pressure drop when thetriple entry upstream manifold was installed. When this difference istaken into consideration and the change of the effective flow area of thesystems is determined at approximately 6000 lb /hr, the single entry up­stream manifold suffered a loss of approximately 24-percent effectivearea, and the triple entry upstream manifold lost approximately 6 per­cent of the effective area.

The afterburner stability limits shown in Fig. 11 are composed ofburner blowout (bottom and left) and the limiting maximum exhaust noz­zle area of 6. 94 sq ft (top). The afterburner was considered to be outwhen burning became so negligible as to produce no effective heatrelease although localized burning continued at the lowest afterburningfuel flows that could be obtained; effective burning then began with anincrease in afterburner fuel flow.

The turbine discharge pressures and afterburner fuel flows atwhich the single entry and the triple entry manifold configurationswould operate are essentially the same at the low ends, within theaccuracy with which the limits could be determined. However, com­bustion could be sustained at lower afterburner fuel flows in the higherturbine discharge pressure region with the single entry manifold con­figuration than with the triple entry manifold configurations. In fact,the single entry manifold configuration lean blowout line is very similarto that obtained during operation with the downstream manifold only.This is caused by the vaporization in the single entry upstream mani­fold at low afterburner flows, which causes a shift of larger percentagesof the flow to the downstream manifold.

Afterburner combustion efficiencies obtained during the test areshown in Fig. 12. At the low values of afterburner fuel-air ratio, oper­ation on the downstream manifold consistently displayed higher combus­tion efficiencies, than with the configurations using the upstream mani­folds, because of the better stability of the flame obtained with the fuelimpingement on the flameholders. The apparent afterburner combustionefficiencies obtained with the single entry manifold configuration weregenerally higher than those obtained with the triple entry manifold con­figuration in the low and high afterburner fuel-air ratio ranges; the tripleentry manifold configuration generally gave indications of efficiencie swhich were higher in the mid-range. It should be pointed out, however ..that the indicated difference in combustion efficiency between the singleentry and triple entry manifold configurations is generally within thescatter of the data values at all but the low afterburner fuel-air ratios.

13

AEDC-TN-59-11

OVERALL ENGINE PERFORMANCE

The engine pumping characteristics and engine generalized airflowvs generalized low pressure compressor rotor speed (Figs. 13 and 14)typify the engine performance results obtained. Generalized adjustednet thrust and generalized specific fuel consumption for nonafterburningoperation at a constant high pressure rotor speed of 7650 rpm and at aconstant Mach number of 1.47 over an altitude range from 43" 000 to63,400 ft are shown in Fig 15. The change in average generalized ad­justed net thrust is shown to decrease only 1.. 6 percent from 43" 000 ftto 63, 400 ft. The average generalized specific fuel consumption overthat range increases 3.8 percent.

The generalized specific fuel consumption vs generalized adjustednet thrust for afterburner operation is shown in Fig. 16. The trend inafterburner combustion efficiency (Fig. 12) at the higher pressure levelsis also shown here; namely" the performance of the single entry andtriple entry manifold configurations is generally the same in the midrange and the single entry manifold configuration indicates higher valuesof afterburner combustion efficiency and lower values of generalizedspecific fuel consumption at the low end. At the lower pressures" thesmaller operating range and fewer data points do not permit the estab­lishment of relationships with the same dependability as at the high pres­sure levels.

SUMMARY OF RESULTS

The results of the performance evaluation of an Iroquois AX-102turbojet engine may be summarized as follows:

1. The overall performance with the single entry manifold configurationis equal to or better than that of the triple entry manifold configu­ration over mo st of the range tested.

2. The rate of afterburner fuel vaporization in the triple entry upstreammanifold was much less than in the single entry upstream manifoldin the lower afterburner fuel flow region; however" this vaporizationin the single entry upstream manifold resulted in an extension of theoperating range of the afterburner in the low afterburner fuel flowregion.

3. Afterburner lean blowout during this test was somewhat of an elusiveparameter since at many conditions spot burning continued althoughthe effective heat release approached zero at the minimum fuel flowpossible.

14

AE DC- TN-59-11

4. A severe surge problem existed during attempts to operate the engineat generalized rotor speeds as low as 7700 rpm.

5. The operational range of the high pressure compressor was improvedby bleeding overboard up to 2. 34 percent of the compressor airflowat the high pressure compressor discharge.

REFEREt-.~CES

1. Mayfield.. Rupert C... "Mechanical Operating Experience With AnIroquois Turbojet Engine In An Altitude Test Chamber. "AEDC-TN-58-101 .. January 1959. (Confidential)

2. Orenda Engines Limited, "Mechanical Description of the IroquoisSeries 2." B4-58, March 1958. (Confidential)

3. Orenda Engines Limited.. "Iroquois 2 Engine Preliminary ModelSpecification." EMS-8 Issue 2. (Confidential)

4. Test Facilities Handbook, (2nd Edition). "Engine Test Facility, Vol. 2."Arnold Engineering Development Center .. January 1959.

5. Povolny.. John, "Use of Choked Nozzle Technique and Exhaust JetDiffuser for Extending Operable Range of Jet-Engine ResearchFacilities." NACA-RM E52E12, July 1952.

6. Britnell, Walter, "Operating Procedure for AEDC Testing of IroquoisEngine AX-I02." July 8, 1958.

15

pA

AEDC- TN-59-11

APPENDIX A

METHODS OF CALCULATION

TEMPERATURE

Total temperatures at stations 2" 3" 4" 7" and 9a were obtained byapplying a recovery factor to the indicated temperature measurementsin the calculation

TiT=-------,,------=-y-l [ Y-l](~)-Y-+ RF 1- (~)-Y-

where RF = O. 85 (stations 2" 3" and 4)and RF = O. 95 (stations 7 and 9a)

AIRFLOW

Airflow at stations In and 2 was calculated from the equation

[ Y-l]y2~~ 1- (~ry

Wa = -~~---=--~-~-y-lVRT (~)-y-

When the venturi was choked" airflow at station In was calculated fromthe expression

Gas flow at stations 5 and 6a was determined by

where 0.02 WaIn represents the manufacturer's approximation of the

amount of cooling airflow extracted at station 4 and returned to the cycleat station 6. Wab1 in the above equation represents the overboard bleed

air extracted at station 4. Overboard bleed air was calculated" using aport flow coefficient of O. 65" to be as follows:

16

AEDC-TN-S9-11

Tests 02-10-1100 through 07-20-1100---0Test s 08- 20 -11 01- - - - - - - - - - - - 0 011 7 W· a1n

Test '09-20-1102 through 17-30-1202---0.0234 WaIn

Gas flow at stations 6 and 7 was determined by

Wg = Wg = Wa + W£ - Wa6 7 in e hI

Gas flow at station 9 was determined by

RATIO OF SPECIFIC HEATS

The ratio of specific heats" y" was assumed to be 1.4 at stations1 and 2. At the remaining engine stations, 3, 4" 5" 6" 7" 8, and 9, thevariation due to gas temperature and composition was considered.

EQUIVALENT FLIGHT CONDITIONS

The free-stream total pressure was obtained by using the standardram recovery propo sed by AlA as in the expression

for Moo> 1.0

and

Poo = P 2

Equivalent ambient temperature was calculated by the expression

t,OO = (P00 ) Y - 1Poo y-

Flight velocity' was calculated by the expression

17

AEDC- TN-59-11

JET NOZZLE INLET TEMPERATURE

The jet nozzle inlet temperature for afterburner operation with thejet nozzle choked was obtained from the expression

[

2

)

1_2_ y=T

_ 2yg Pg ( Y + 1 AT.. - R (y + 1) w. • 1. • fJ

For non-afterburner operation T 10 was assumed equal to T 7.

EFFECTIVE NOZZLE AREA

Non-Afterburning

The effective nozzle area during non-afterburning operation wasobtained from the following expre ssions

Nozzle Choked

Nozzle Unchoked

~.r _(.~)~]\l y - 1 L1P g

Afterburning

The effective nozzle area during afterburning operation was obtainedfrom the following equation

The exhaust nozzle coefficients (Cd) that were used are shown inFig. 17.

18

AEDC-TN-59-1l

FUEL-AIR RATIO

Engine fuel-air ratio was defined as

Wffe

= __e_

was

Afterburner fuel-air ratio was defined by the expression

where 14.62 was assumed equal to stoichiometric air-fuel ratio for thefuel used.

Total fuel-air ratio was obtained by the expression

THRUST

Sea Ie Thrust

Scale jet thrust at the desired flight condition was calculated fromthe equation

Net Thrust

Net thrust was then determined by

F n d'Sa ]

19

AEDC· TN·59·11

Calculated Thrust

Ideal thrust for a converging nozzle was calculated from

y+l

VT (Y~1)2(Y-l)

Ideal thrust resulting from isentropic expansion to test chamberpressure was determined from the expression

F.. ~ w rP:T:J2L [l-(~)Y;1 ]lIs gg \I g' y - 1 P 9

Thrust Coefficient

The convergent nozzle thrust coefficient was defined as

F·CF. = _l_s_

len F ien

and the isentropic thrust coefficient was defined as

COMBUSTION EFFICIENCY

The combustion efficiency in the main engine combustion chamberwas determined from the following equation

1Jb

where the quantity 59.62 Btu/lb fuel accounts for the difference betweenthe enthalpy of carbon dioxide and water vapor formed during combustionand the enthalpy of the oxygen removed from the air by their formationin the temperature range from 4000 R (base of enthalpy equations) to 5400 R(base temperature of determination of hL ).

The combustion efficiency of the afterburner was calculated from

1J ab =

20

AEDC-TN-59-11

COMPRESSOR EFFICIENCY

The overall compressor efficiency was defined as'

"'c =

where Ha is the theoretical enthalpy that would be obtained at station4is

4 if isentropic compression had been experienced.

The compressor efficiency of the low pressure compressor wasobtained from

The high pressure compressor efficiency was calculated from

TURBINE EFFICIENCY

The turbine efficiency was defined as

where Y tu was defined as being

(H gs - H g2 )

( T s - T 7 )

(H gs - H g2)

(Ts - T 7 )

_R__

778.3

21

AE DC-TN-59-11

TAILPIPE EFFICIENCY

The tailpipe efficiency was defined as

1Jtp

Y7 - 1

(~) Y7 - I

SPECIFIC FUEL COMSUMPTION

The specific fuel consumption was obtained from the followingexpressions

S F C

REYNOLDS NUMBER INDEX

Wf + Wf be a

Fns d'a J

The Reynold number index was calculated from the equation

ReI = _--,°3::...-_cP2 V8a

TURBINE INLET PRESSURE

o~(T2+216)

735 822

The turbine inlet pressure was obtained from the expression

which was obtained from the engine manufacturer.

TURBINE INLET TEMPERATURE

The turbine inlet temperature was calculated by the iteration processfrom the following equation

0.2318 T s + 0.052 X 10-4 Ts 2 + 0.2388 X 10-aTs3

- 93.705

I + f e

+ f e [0.2655 T s + 1.8632 X 10- 4 T/ - 2.2117 X 10-a T S3

- 134.60]

22

AEDC- TN-59-11

when T5 is equal or less than 1700° R.

When T5 is greater than 1700° R the following equation was used

-4 2 -8 3H _ 0.2214 T s + 0.176 x 10 T s - 0.1258 x 10 T s - 93.705gs - 1 + fe

[ -4 2 1 - 8 3_+ f e 0.3397 T s + 1.359 x 10 T s - 0.9861 x 0 T s

where

176.15]

Hgs

PRESSURE LOSS DUE TO HEAT ADDITION

The tailpipe pressure loss due to heat addition was determinedfrom the expression

COMBUSTION CHAMBER VELOCITY

Y8 - 12

Y8M 2) Ys - 1

8 a

The combustion chamber velocity was calculated from the expression

V y-l1 + 2

and Mcc was determined from the equation

~}g MR cc

where Acc = 7.674 ft sq.

23

FLOW COEFFICIENT

The flow coefficients were determined by the expression

AEDC- TN-59-11

AP

24

AEDC-TN-59-11

..cu

..cECJ

.cU...

IIIcu

l-e

co+:

CJ

.2III

-=~..o0i:Q.

cu05

I:J)c

WNo..­

I

><«

.~U-

25

t\:)~

Fig. 2 Internal View of Iroquois Afterburner

»mo\1-IZtn'P

AEDC- TN-59-11

35 - 0.031 "DIA. HOLESEQUALLY SPACED

70- 0.027" DIA. HOLES

EQUA",LY SPACED

MANIFOLD ------::~

FUEL FLOW

A. SINGLE ENTRY MANIFOLD

~~:. g:~~~','. ~',~'. ~gt~~ a -------EQUALL Y SPACED

35-0.031" DIA. HOLES a35-0.187" DIA. HOLES

EQUALLY SPACED FLOW

B. TRIPLE ENTRY MANIFOLD

FUEL rtJEL FLOW

ORIFICE SIZEO.

I 0.020

2 0.018FIVE EACH

3 0.022

4 0.022

5 0.022

ALL VIEWS LOOKINGDOWNSTREAM

NO.1 a 4

GAS FLOW -c=-­COMPOSITE CROSS SECTION

C. DOWNSTREAM MANIFOLD

Fig. 3 Afterburner Fuel Manifolds

27

Nco

'2 I I ! I \90INCHES

Fig. 4 Installation of Iroquois in Altitude Test Chamber

>mo()

~Z0,'f>----

PLENUM 1Ijo 1

LABYRINTH

I ~ {SEAL VENTURI1 .. (THROAT

' .:.....H--

1"" -II_~ L 51L.::l-- 114.0 ==j~-~---

DOWNSTREAM AFTERBURNERFUEL MANIFOLD

. l~~ 3L~

9 I I I I 19°INCHES

I:\:)

CO

Fig. 5 Test Article Station Locations

»moQ-fZIn'P......

AEDC-TN-59-1l

2'T(f-.....-

a. STATION In, VENTURI THROAT, 137 IN.UPSTREAM OF COMPRESSOR FRONT FACE

c. STATION 3, LOW PRESSURE COMPRESSORDISCHARGE 1.38 IN. DOWNSTREAM OFSTATOR EXIT

b. STATION 2, COMPRESSOR INLET,. 5.25 INUPSTREAM OF COMPRESSOR FRONT FACE

0"

d. STATION 4, HIGH PRESSURE COMPRESSORDISCHARGE O. 65 IN. DOWNSTREAM OFSTATOR EXIT

o TOTAL PRESSURE

• STATIC PRESSURE

)( THERMOCOUPLE

Fig. 6 Instrumentation Location Details (looking upstream)

30

AE DC-TN-S9-11

" _~_B5C

e. STATION 7. TURBINE DISCHARGE(SINGLE POINT THERMOCOUPLESELECTRICALLY AVERAGED) 0.96 IN.DOWNST REAM OF TURBINE DISCHARGETRAILING EDGE AT MEAN CHORD

g. STATION 9. EXHAUST NOZZLE INLET55.4 IN. DOWNSTREAM OF AFTERBURNERFRONT FLANGE

f. STATION 9a. AFTERBURNER. 38 IN.DOWNSTREAM OF AFTERBURNERFRONT FLANGE

h. STATION lOs. EXHAUST NOZZLEAMBIENT, OUTER SURFACE OFNOZZLE FLAPS IN PLANE OFNOZZLE EXIT

o TOTAL PRESSURE

• STATIC PRESSURE

X THERMOCOUPLE

Fig. 6 Concluded

31

80

90 ' 'Hi:tt+tf 11III1 ATI Ii Iii I III I I II I I I I III I1III11I1 I t-tt+ttt1

o 'Simulated flight conditionsat which engine evaluationdata wer e obtained

• Simulated flight conditionsat which afterburner datawere obtained

70

60

('I")

t0...... 50

><+.l......

..Q)

"C:::s 40+.l."+.l......«

30

20

10

AEDC- TN-S9-11

..l.

i

Pr e1iminaryModel Spec. Ref. 5

o0.0 0.4 0.8 1.2

Mach Number

2.0 2.4

Fig. 7 Iroquois Series 2 Operational Envelope

32

AEDC-TN-59-11

15

14

13

12('t')

I0

111"""4

>l

"'"~ 10----,.Q1"""4

,.Q9n5....

~~

8~0

1"""4

~

1"""4 7Q):;j

~

"'" 6Q)

~J.l:;j

,.Q

"'" 5Q)~

<4

3

2

1

00 1

--Single -~ntry ~~n1~~iIAII:~~uration'

OTriple entry manifold configuration ­

ODownstream manifold only

N2

= 7650 rpm

T 2 - 560 0 R

2 3 4 5 6 7 8 9 10

Engine Inlet Total Pressure, P 2 J psfax 10- 2

Fig. 8 Afterburner Test Data Obtained

11

I .,

o

12

33

7.0

o 0 percent station 4 overboard bleedo 1. 17 percent station 4 overboard bleedo Z.34 percent station 4 overboard bleed

guide vane at 00 position-t-l-

~-d.. r,ll

6.5 T Z =41So R

ReI = 0.37

.t:tJl{I)

....,{I)

~ 5.0

~J::tl

1. 17% overboard bleedott

oMo~

<~CD

.:;: 5. 5 2. 34%

CD overboard'il • bleedNoZ

4.5-:

.... -t,

ff

i.L+-H ;

High Pressure Compressor Rotor Speed, NZ' rpm

Fig. 9 High Pressure Compressor Surge Limitw~

4.06000 6500 7000

i-Itt-~

7500 8000 »mo()

~z0,'P

AEDC- TN-S9-1l

I=i:-±::E -+-1-

25H-

MIo.-4

=f -20 I

ITf-f- --

Ft:15 I

5

H-'-,-l-

'UU

Total Fuel Flow

o Single entrymanifold config.uration

o Triple entrymanifold configuration

6. Downstream manifoldonly

3

~=p-rl--',-,-

4

2

I 1

1-+--+--++++++++++H-+t-HH-tti-tti1

1--W-W--l-W-I-W-l--I+~--W-l--t-\-I-+-\-t-\-l-l+l+1-f-1+H-l+H-H+H-H++++t+H-t+t+ttttt+tftttttt-t+t-1itHHiii1iiiit-rH-t-mHilH1iitt'HttlitttttmtmtL1

1 2 3 4 5 6 7 8 9 10 15 20 25 30

Afterburner Fuel Manifold Pressure Differential

p~ -P9 , psid x 10- 3

Fig. 10 Afterburner Fuel Manifold Flow Characteristics

35

12

14--- Constant Exhaust Nozzle Area

-- ----Afterburner Blowout Limit- ... - Afterburner Stability Limit

IIII · Data Manually Takeno Steady State DataN2 7650 rpm

T2

-560oR

AEDC-TN-59-11

6.94. (1000)

~

I0 -..--4 N~ 10 60 60 .SJ..4 - - (950) ;;..c:

--- .::::~.....

':..... 0~ ..--4I'd <~

~ 8 ..I'd

'i Q)

6.25 J..40(900) ~.....

~ .......... NQ) N~

6 0~ Z-J..4 ~

Q)5.90 til

~ ~J..4 (850) I'd~ ..cl~ ~

J..4 riIQ)~

4< 5.56(800)

..L

2

o4 8 12 16 20

Turbine Discharge Pressure, P7

, psfax 10- 2

a. Single Entry Upstream Manifold Configuration

Fig. 11 Afterburner Stability Limits

24

36

14---Constant Exhaust Nozzle Area-----Afterburner Blowout Limit- ......Afterburner Stability Limit

• Data Manually Takeno Steady State Data

N Z = 76~0 rpm

T 2 = 5600

R

AEDC-TN-59-11

6.94(1000)

M -NI .S0...-t -~ 10 N

6.60 ¢::J-4 (950)..Q ....~-.. 0..0 ...-t...-t <..

..0 ...cO cO

'+-4 8 Q)

~ J-4

<'i -t- -I Q)

0I-r

6.25 ...-t...-t N

~ (900) N0

...-t ZQ):;j 6 ~

~ rn:;j

J-4 cOQ) 5090 ~r:lJ-4 (850) ~:;j..0

J-4Q)

5~56~ 4"+-l

< .I (800)

2

o4 8 12 16 20

Turbine Discharge Pressure, P7

, psfax 10- 2

b. Triple Entry Upstream Manifold Configuration

Fig. 11 Continued

24

37

AEDC- TN-S9-11

+t ..

14 t++oIl4+++++-I Constant Exhaust Nozzle Area111--- Afterburner Blowout Limit

~ill-···"AfterburnerStability Limit

.. Data Manually Taken~ Steady State Data

N2

7650 rpm

T 2 5600

R

10

6.25(900)

. -I.

5.90(850)

-t 5.56(800)

-Nt=:.....-

.....o

6.60 .....(950) ~

nlQ)

~

<Q).....NNoZ

T-+ --~~, 1-

_ LL

-+-

I

1-

4 +

!::l

f .J

~=-'

21 -'r.:

MIo.....

-I- .± -+ ,-c'

:1: _ ~: .. ~.

8 12 16 20

Turbine Discharge Pressure, P7

, psfax 10-2

o4

+-tt- ~+H- H­-tt j~~: r+ . H··t

.,

-1+,j I . - t-+

24

c. Downstream Manifold Only

Fig. 11 Concluded

38

AEDC-TN-59-11

, ::t .

. '1

1:'iEt . tt·d+· '''IT

c. P 7 - 1450 psfa tI~fB,... , r" 'f

+H·!;4

'1

t"l'

+1· .

H r~h+ 8:. ..:

+1 H--

.. t 1. +.+,=! 11 . Jt

:rt f' 1 U=t ~¢ 3t'!ht ±<=it:;

:•• ~H."'I.h!"H++''++++ d. P7 ~f1f~~j~ffc~~11mtIT-r1*Ilm.n:;

i :1

u·

=t~ j; 1 '

~1I1+

tr t

Wij uI~~ :t !11·;t 1 j1iLjfIih I; I:; It f\l£~L i1m;; iiHnn ifni]) :;;;l1t iT if1'+ t!1; rt b. P7 - 1800 psfa IL:.i+~,ttj:;tlt;:j:I:j

I!H i J+. e'i:: ~:i t i 11 :t 11 Ii t h ~tTi=

t

80

80

70

80

70

70

90

90

60

100

d,

't"

.;.,-L +'

Itr ' e' n t. "" Ii't' 't,, ,". ".' "+,' .,0·1·

0.01 0.02 0.03 0.04

60

100

90

80

70

60o

'tt j. i ti:

; .. -PI.,Configuration " I

o Single Entry Manifold =:

o Triple Entry Manifold ~;....

~ Downstream Manifold mf

f"i 1+"·,; H t ;;;!;;; 11:

P i; e. P 7 = 850 psfa;;~':HH~

. .. f;i~; ~1~!+rH;i+-:lr1f-=ln~'lr! ,•. .:i: IjH.;Ii ,I; t+l'i:;n;+'I1:1i:Hi' ;lj; !.';j'L

0.05 0.06

Afterburner Fuel-Air Ratio, fab

Fig. 12 Afterburner Combustion Efficiency

39

1.5

AEDC.TN-59-11

..........I'llI-lQ)

b 1.0 o 00/0 station 4overboard bleed

() 1. 17% station 4overboard bleed·

• 2. 34% station 4overboard bleed

ReI=0.37

T 2 = 4150 R

AID. =4.5 .. 5.28 ft21

.~. 0 2.0

Overall Engine Temperature Ratio, T7/ T 2

Fig. 13 Engine Pumping Characteristics

3.0

40

AEDC- TN-59-11

240

4.68

4.865.07

220 5.28

ReI =O. 37

T2 - 415 0 R

o Al O. =4. 5 ftl1

CJ AlO i =4.68 ft 2

<> AlO. - 4.86 ft21 2

L:i. AlO i =5.07 ft

'V AIO i = 5.28 ft 2

50004600

Open symbols indicate00/0 station 4 overboardbleed

Half closed symbolsindicate 1. 170/0 station 4overboard bleed'

Closed symbols indicate2. 34% station 4 over ­board bleed

420038003400100

3000

120

140

160

180

Generalized Low Pressure Compressor

Rotor Speed, N 1/rei. I rpm

Fig. 14 Generalized Compressor Airflow

41

AEDC· TN-S9-11

lOx 10 3

..a........

tN1.0-..... 9.~

"0rd

til$:l

r.z.t

8

a. Generalized Adjusted Net Thrust

o P Z 1152 ps£a; Alt 43, 000 £t

D P z 907 ps£a; Alt 48, 100 £t

6. P2

734 ps£a; Alt 52,500 £t

<> p 2 57 6 p s£a; Alt 57, 400 £t

'\l P 2 432 ps£a; Alt 63, 400 £t

oT 2 =560 R

N 2 - 7650 rpm

Station 4 Over­board Bleed 2. 340/0

j2.0til

$:lr.z.t-.......a.....-.....J.l

~ 1.5..a.....

b. Generalized Specific Fuel Consumption·

65 x 106045 50 55

Altitude, £t

Fig. 15 Non-afterburning Engine Performance at Mach Number 1.47 at Altitude

~-.....Ur.z.t 1.0U) 40

. 42

r-l(1)

::1 1.4I'.z.tu

-.-j

'+-i 2.l• .-j

u(1)

0..(J) 2.0"0(1)N

-.-j

r-ltilH(1)

J::(1)

()

l.l

2.0

AEDC· TN·59·11

Configuration

o Single Entry Manifoldo Triple Entry Manifold6. Downstream Manifold

T 2 ::: 5600 R

N 2 ::: 7650 rpm; P 4/P7 ::: 3.2

Station 4 Overboard Bleed 2"=-.~:::-,-v,"":"~!":~'-:-4~'-'~}~+-=~:~~F:_.-'t-:-:.~I-'::-~:"+'-+"~+fm::~J7: :;)::~:F:F~: :TiL::;

9 10 11 Il 13 14 15 16 x 10 3

Generalized Adjusted Net Thrust Fnsad/ °2 lb

Fig. 16 Afterburning Engine Performance at Mach Number 1.47 at Altitude

43

1. 00""""" '_fttI+H++lllllllllll! III i i III ftffi-f-tttj

"0U

.+..~

~Q)

-r-! .98_ P 9/ Po Range: 2.5 to 5.9U-r-!..........Q)

0I. ,

UQ)

heJ-tcd .96..c:u00

-r-!

QQ)

...... +-,NN

. +f0 II..... .94 II~ r~ ,I I I I., f-'-'-~

00

=' f-f+Rcd..c:X

f;J:1

~

~

4.0 5.0 6.0

Exhaust Nozzle Area. A IO _' ft21

Fig. 17 Exhaust Nozzle Discharge Coefficient

7.0

>moQ-I~OJ'P....