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RESEARCH MEMORANDUM FOR REFERErICE

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EFFECTS OF ALTITUDE ON TURBOJET ENGINE PERFORMANCE

By William A~' Flenring

Lewis Flight Propulsion Laboratory Cleveland. Ohio

CLASSIFICATION CHANGED

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON December 21, 1951

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NATIONAL .ADVISORY COMwlli'I'n FOR AERONAUTICS

BE3EAROlI MHM:>RAND'UM

EF.F'ECTS OF .ALTITUDE ON TORBO.:JEr ENGINE PEBFOBMANCE

By William. A. Fleming

Component and over-all performance characteristics of several turbojet engines investigated at the N'.!\CA Lewis laboratory during the past several years are summarized in this report to indicate the effects of altitude on turbojet engine performance. Data presented herein were obtained over a range of altitude flight conditions in the altitude wind tunnel and aD. altitude chamber, and alSo over a range of inlet conditions in compressor component installations.

A summarization of data for several engines shows that failure of turbojet engine performance to generalize for all altitudes is primarily due to a decrease in ccmpressor ef'ficiency and corrected air flow with reduced Reynolds number index and a reduction in combustion efficiency with increased altitude. Data also show that, although engines of dif­ferent design may have equal thrusts at sea level, the thrust can dif'fer by as IDllch as 26 percent at an altitude of 45,000 feet. These differ­ences, which are introduced as a corrected engine speed effect, are primari~ due to differences in compresso~ air flow and efficiency char­acteristics from. one engine to another. The additional effects of Reynolds number on the performance variations with altitude from. one engine to another are of much smaller magnitude than the corrected engine speed effect.

INTRODUCTION

The effect of altitude on turbojet engine performance can be approx­imated by the application of correction factors to the performance vari­ables obtained at sea level. In the same manner, data obtained at sev­eral aititudes may be generalized or corrected to a ref'erence altitude, usually sea level. Performance variables thus treated are referred to as "corrected" or "generalized" performance, examples of w.b.ich are corrected engin.e speed NI -.,f8 and corrected net thrust Fn/5. The purpose of the factor 8 is to correct for the ef'fect of ambient pres­sure on the density of the engine air, and. the purpose of the factor e is to correct for the effect of temperature on the Mach numbers through­out the engine for each rotor speed. Definitions of the factors 8 and e, together with a complete list of corrected perf'ormance variables, are

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given in the ~ppend1x. One derivation of' these corrected perf'ormance variables which utilizes the concept o~flow similarity is presented in reference 1.

In order to indicate deviations in peri'ormance variati.ons with altitude among a number of turbojet engines and the causes of' these deviations, thereby to provide a better understanding of the effect of' altitude on turbojet engin~ perf'ormance, dataobtained~om investiga­tions of several' ,axial- and centrifugal-flow engines at the NACA Lewis laboratory during the last several years are summarized in this report. Discrepancies between actual engine perf'ormance at altitude and altitude perf'ormance predicted fr~ sea-level data are crit~cally- examined ~or' the various engines and. the reasons for these discrepancies are traced. t.o particular engine components. In addition~- the·-·proD.ciiiri.6ed effect of:" differences in COmpressor characteristics ~Om one engine to an.other on the rate. of' thrust decay or variation .of maximum corrected engine thrust with altitude is also examined.. Data thus presented. herein include com­ponent and over-all engine perf'ormance obtained in the altitude wind tunnel and analtitua.e test ch~ber, and. cqmpress.or perf'ormance obtained. in compressor installations.

APP.ARATOS

Altitude wind tunne1.- The altitude wind tunnel in ~hich some o~ the engines were investigated is a closed-circuit return-type tunnel circular in cross secti.on·-with a test section 20 ~eet in diameter and 40 feet long. Auxiliary tunnel equipment, including exhausters, refrigerati.on system, and. drive mot.or and fan, are used to s:1lnulate alti­tude flight conditione. Engines being investigated. are generally mounted in a fuselage or nacelle (~ig. 1), which is connected to a six-component scale system by means o~ trunni.ons at either side of' the teet section.

Because it is desirable to simulate flight Mach numbers greater than those obtainable with the tunnel fan, dry refrigerated air ~cm the tunnel make-up air system is introd.uced through a duct t.o the engine inlet. 'l!b.is air is throttled ~rom. approximately sea-level pressure to the desired total pressure at the engine inlet while the static pressure in the test section of' the tunnel is maintained. at the des,ired. altitude condition. With this .operating technique, the test-eection pressure can be reduced. to corresp.ond to altitudes as high as 50,000 ~eet while the simulated flight Mach number is varied from. zero t.o approximately 1.0. BY' the use of either electric heaters or refrigeration coils installed in the .make-up air system, the engine-inlet temperature can be varied from 150.0 to _300 F.

Altitude chamber. - The altitude chamber in which some .of the engines were investigated is 10 feet in diameter and 60 feet long. For centrUugal-f'low-engine ins~allatione, the chamber is divided int.o two

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sections, an inlet section and an exhaust section, separated by a steel bulkhead located at the rear of the engine. For axial-f'low-engine instal­lations, an additional bulkhead is installed at the engine inlet, 'Which isolates the engine section f'rom the inlet section (f'ig. 2). With this type of' installation, small amounts of air are bled through a valve in the f'orward bulkhead and exhausted. tllrough an opening in the rear bulk­head to provide circulation in the engine section. ~e engines are m01mted on the trunnions of a thrust-measuring bed with the tail pipe of' the engines extending through the rear bulkhead into the exhaust section. Freedom of' movement for the engine at the bulkheads is provided in both axial and. radial directions.

Air at ram pressure is supplied to the inlet section of' the chamber through a supply line f'rom the laboratory air system. The temperature of' this air can be varied f'rom aSC to -500 F. In order to insure a unif'orm. velocity profile at the engine inlet, the air is passed through a set of' honeycomb vanes installed in the upstream end of the chamber. The hot gases discharged f'rom the engine are removed. f'rom the exhaust section of the altitude chamber through a dif'f'using elbow and passed through two dry-type coolers before entering the laboratory exhaust system. The pressure in the exhaust section can be reduced to corres­pond to altitudes as high as 65,000 feet.

Compressor component installations. - In the compressor component installations the compressor, which ~ driven by a variable-.frequency induction motor through a gE!ar box, was mounted. downstream of a stag­nation chamber (fig. 3). Screens were fitted in.:to the tank near the midsection to remove any foreign particles and to insure smooth f'low. The inlet air_ passed through a submergedad.Justable orifice in the inlet ducting, intq the stagnation chamber, and then into the compressor. The air leaving the compressor paSsed. through a collecting chamber and into the laboratory exhaust system. CanJl)ressor-in1et and -outlet pres­sures were regulated by butterf'ly throttle valves.

Inlet temperature and inlet pressure can be varied over the same range as in the altitude chamber. The inlet ducting, stagnation chamber, and. the compressor were insulated to minimize heat transfer between the working fluid and room air.

Engines and compressors. - Data obtained with seven engines and two compressors are included in this report. The engines were designated as A to G; engines A to D were axia1-f'low engines and eng:tnes E to G were centrif'uga1-f1ow engines. The two compressors operated in compressor component installations were silIl.i1ar to thee e f'or engines C and. F and are designated. as such, but were in stages of' development different from the ones uSed in the engines for the complete engine evaluations. Com­pressor a had only 10 stages, whereas the com:pressor in engine C had 11 stages. Characteristics of the engine com:ponents are summarized in the f'ollowing table:

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Eng:ine Com;pressOT Combustor Turbine Air flow Compressor at rated pressure sea-level ratio conditions at rated

(lb/sec) sea-level conditions

A l2-stage 8 tubular type Single stage 94 5.1 axial through flow

B ll-stage 8 tubular type Single _stage 85 4.8 axial through flow

C ll-stage doub le annular Two-stage _.

58 3.8 axial type through flow

n ll-stage 8 tubular type Single stage 73 4.0 axial through flow . -_. ---

E double entry 14 tubular type Single stage 78 4.0 centrifugal through flow

F double entry 14 tubular type Single stage 90 4.6-centrifugal through flow

G single entry 9 tubular type Single stage 91 4.4 centrifugal through flow

- - ... _.- .. _ .. ..- -

INSTRUMENTATION

For the full-scale eng:ine investigations, total and static pressures and indicated temperature were meBBured. at the :inlet and outlet of each component, except at the turbine inle:t, Where generally only the total press:ure was measured. Sketches of' survey rakes typical of those installed at f!everal circumf'erential positions at each station in the engines are shown in figure 4. Engine -air f'low was determined from the compressor-inlet instrumentation on the axial:flow engines and with the tail-pipe-outiet instrumentation on the centrifugal-flow engines.

For the compressor component installations, pressure and tempera­ture were measured in the stagnation chamber to determine the state of the inlet air. Compressor-outlet measurements were made with lnstru"; mentation similar to that used in the full-scale engines. Compressor air f'low was measured bya submerged adjustable orifice located_in a straight section of .the inlet ducting upstream. of the stagnation

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PROCEDURE

When the engine perf'ormance data were being obtained, the engine­inlet total pressure and temperature and the exhaust-section or tunnel­test-section static pressure were set to correspond to the desired flight conditions for an NACA standard atmosphere, assuming complete free-stream ram pressure recovery at the engine inlet. The engine speed was then varied from the minimum to maximum value and both ,over-all and component performance data were obtained simultaneously. At each rotor speed inves­tigated in the compressor component installation, the inlet-air tempera­ture and pressure were ma.in.tained constant and the air f'low was varied from the maximum obtainable with the laboratory air system. to the point Just preceding surge. Surging was detected audibly and by observing fluctuations on the p~essure measuring manometers.

Methods of calculating each of the performance variables presented herein are discussed in the appendix.

RESULTS AND DISCUSSION

Effects of Altitude on Generalized Performance

When turbojet engine performance obtained at several altitudes was generalized to standard sea-level conditions by use of the correction f'actors 6 and e, it has been found that data obtained at high altitudes often deviate :from. those obtained at lower altitudes. An example of the types of deviations encountered with generalized engine performance is shown in f'igures 5 to 8, where perf'ormance data obtained at several alti­tudes with an axial- and a centrifugal-flow engine, engines B and F, respectively, (references 2 and 3) are corrected to standard sea-level conditions. - These data, which are typical of most current axial- and centrifugal-now engine performance, show that application of the temper­ature and pressure correction factors e and 6 failed to generalize each variable to ,a single curve. As indicated 'in these figures, the effects of al ti tude on the performance variables often became apparent only above altitudes of' 30,000 to 40,000 feet. It is these secondary effects at high altitudes which will be of' primary concern in the suc­ceeding discussion.

With respect to the effects of altitude variations at a given,cor­rected engine speed, an increase in altitude raised the corrected net

... thrust of the axial-flow engine and reduced the corrected thrust of the centrifugal-flow engine (fig. 5). For the other performance variables presented, an increase in altitude resulted in a decrease in corrected air flow (fig. 6), an increase in corrected specific fuel consumption (fig. 7), and an increase in corrected turbine-outlet temperature (fig. 8) for both engines. ' For' a full understanding of the basic causes of these

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altitude effects, it-becomes necessary to trace the effects to their origin by examining the altitude perforIn@c~ 0.1'. each eng~e component.

Effect of Inlet Conditions on Component Performance

Generalized data are presented showing the performance of each com­ponent over a range of inlet conditions in order to isolate the components affected by altitude. For each component, data for a number of engines were selected to show that the trends are typical of current engines.

Compressor. - The typical effect of altitude on compressor perform­ance characteristics is illustrated in figure 9, in which performance maps are presented at two different compressor-inlet total pressures for the axial and centrifugal compressors of engineEl C and F operated in compressor component installations.· The effect of inlet pressure on the variation of ,compressor pressure ratio with corrected air flow is shown for both com'pressors~t three corrected compressor speeds with constant-efficiency ·contouTssuperlmPosed. The reduction in inlet pressure shifted the constant compressor speed lines and efficiency contoursasshQwn in~ the figur~_."",:~1JnP':lblished data further indicate that the peak efficiency of these com:pi-essors---decreased .. as inlet pressure was reduced. . - - .'l--"-.-'-~' _ ... __ ,

In order to indicate further effects of inlet pressure and to show the effect of inlet ~emperature, variations of compressor efficiency and corrected air flow with inlet total pressure are shown in figure lO for engine A (reference 4) at a constant corrected compressor ~eed, pressure ratio, and two inlet_total temperatures. Lowering eitheF the inlet pres­sure or raising the inlet temperature resulted in reductions in both -efficiency and corrected air flow.

It has been found that these effects of altitude or inlet pressure and temperature on compressor performance can be traced to an effect of Reynolds number on flow through compressors, as discussed in reference- 4. These effects on performance B.J:'e introd:t:i.ced· bi the- siil£t~ in- -floW' trans-­ition region on compressor blades with variation in Reynolds number in a similar manner to the well established shift in flaw transition region on isolated airfoils. The effect of this shift in flow tranSition region wi th reduced Reynolds number was to reduce the air flow handling capacity and efficiency at a-given corrected speed and pressure ratio in the manner illustrated in figures 9 and lO for reductions in inlet pressure and increases in inlet temperature. _

The association of Reynolds number with these effects is further illustrated by a correlation of the compressor performance of engine A with Reynolds mnnber index, as shown in figure II for three corrected compressor speeds. For convenience, the parameter Reynolds number index

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is used in place of' Reynolds number because of' the difficulty in select­ing corresponding dimensj,ons on which to base Reynolds number for use in cOlIlparing axial- and centrifugal-flow cOIqpressors. The relation between Reynolds number index and Reynolds number, as developed in reference 5, is included in the appendix. Data for each speed were obtained at two widely different values of inlet temperature; inlet pressures were selec­ted for each temperature such that Reynolds number index remained con­stant. Although. the independent variations of inlet temperature and pressure shown in figure lO had a marked effect on performance, selecting . the combination of inlet pressure and temperature so as to maintain Reynolds number index constant resulted in satisfactory correlation of the data. This correlation is typical of that illustrated by additional data included in reference 4 for a wide range of inlet pressures.

Variation of corrected air flow and efficiency with Reynolds number index is shown in figure l2 for five cOIqpressors (A, reference 6; C, ref­erence 7; D,reference 8; E, reference 9; and F) operating -at rated corrected com;pressor speed and constant pressure ratio. The decrease in corrected air, flow alid efficiency was evident for each compressor, although the effect was more pronounced with some cOIqpressors than with others. Such a pronounced Reynolds number effect on the performance of current turbojet compressors is an im;portant factor contributing to the secondary

1 effects of altitude on over-all engine performance. This relation will be discussed more fully in a later section.

Turbines. - Any altitude effects on turbine performance might also be expected to originate from variations in Reynolds number. Because the pressure and temperature, and consequently Reynolds-number index, at the inlet of a turbine installed in an engine vary with engine speed at a given flight condi tiop., it becomes most .expedient to com;pare the data on the basis of engine-inlet total pressure rather than turbine Reynolds number index.

The effeet of engine-inlet total pressure on the perfo~ce of the turbines in four engines (A, reference lOj Bj D, reference ll; and F) is shown in figures l3 and l4. In no case was there any effect of inlet pressure on the corrected turbine gas flow. Turbine efficiencies were also unaffected except that with very low engine-inlet pressures, l8l and 262 pounds per square foot, the efficiencies of turbines B and F

were reduced by l~ to 3 percent. At these very low inlet pressures,

which correspond to turbine Reynolds number indices in the region 'of 0.08 to 0~l3, Reynolds number effects are becoming apparent in the tur­bine. Because these effects were ~parent only at engine-inlet pressures which would be encountered at altitudes above about 50,000 feet, the per­formance of current engines operating u;p to this altitude should be essentially unaffected by Reynolds number effects on turbine performance •

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Combustor. - Because the exact meChanimn of combustion in a turbojet combustor is not yet fuJ..1.Y understood, no technique for generalizing com­bustor performance is universally accepted. Attempts have been made to generalize the data by plotting combustor performance, that is combustion

P2 - P3 efficiency ~ and total-pressure-loss ratio P ,as functions of 2

corrected engine speed N/ -/82, as in reference 12 ° One experimentor (reference 13) developed a correlation of combustor performance in which

P2T2 combustion efficiency was a f'unction only of y-o Others working with

2 combustor performance have empJ.oyed still different techniques which were basically similar to these. Thus far, however, no technique has been found which will satisfactorily generalize the combustor performance of a large number of different engines.

Because no satisfactory generalizing techn:1,.ques are known, the com­bustor data herein are presented to show the effect of altitude or pres­sure on performance, with no attempt to generalize the data. Combustion

P2 - P3 efficiency and combustor total pressure loss ratio are shown P2

as fUnctions of engine-inlet total pressure at three corrected engine speeds in figures.1S .and. 16, reSpectiv~ly,for ;four engines (A, refer­ence 12; B; ·C; rei'erence 14J aiLd F). As engine-inlet pressure is varied at a constant corrected· engine speed, the combustor-inlet corrected velocity V2/,;e- and·corrected te~eratuXe T2/erematn essentially constant; any small variations have a dec~ed secondary effect on com­bustor performance.

The variation of combustion efficiency with engine inlet total pres­sure (fig. lS) illustrates. the characteristic decrease in efficiency with reductions in both engine inlet pressure and corrected engine speed. As shown, the effects of both pressure and corr~cted speed differ gr~atly from one engine to another; little effect on the combustion efficiency was noted in engine A as cOIllpared )lith the other engine.s. The decided reductions in efficiency at reduced corrected speeds and inlet pressures significantly increase the specific fuel consumption of engines at high altitudes. Also, the. large differences ihthese effects from one engine to another account for similar differences in specific fuel consumption among engine.s.at high altitudes, although the sea-level values may be the same.

There was little effect of inlet pressure on the combustor total­pressure-loss ratio (fig. 16). The total~pressure-loss ratio decreased slightly as inlet pressure was reduced, except for engine F, which exhibited the opposite tre~ •. Thes-= :trendl3 . .t .w~ch .. ine~g·ll.case amgun:t~Q. to -less than a l-percent rea..uction .in total pressure, are considered of

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secondary importance, since a combustor total pressure loss of 1 percent will reduce the thrust and increase specific fuel consumption by about 1 percent. Reductions in corrected speed from the rated ~e to 80 per­cent of rated speed increased the total pressure loss by as much as 2 percent in some cases, although the pressure-loss ratio of engine B was reduced slightly. These trends with corrected engine speed similarly influence the variation of specific fuel consumption with corrected speed.

Exhaust system. - Changes in the swirl a.ngl.e of the flow leaving the turbine with engine operating condition or change~ in Reynolds number affecting the f'low separation from the walls of the tail pipe can affect the total pressure losses through the exhaust system. Since these total pressure losses are reflected by reductions in jet thrust, the pressure loss across the exhaust system can be represented in terms of the ratio of measured to theoretical jet thrust, where the measured thrust is pb~ined from balance-system measurements and the theoretical thrust is based on the total pressure, total temperature, and gas flow entering the exhaust system. In order to indicate whether engine operating con-di tions or Reynolds number affected the flow through the exhaust system of' the engine, the ratio of measured to theoretical Jet thrust was plotted as a function of' pressure ratio across the exhaust system for two engines (A and C) at several ambient pressures (fig. 17). These data cover a wide range of engine operating conditions, including vari­ations in both altitude and engine speed. Although the scatter of the data was fairly large, there were no evident trends of jet thrust ratio with ambient pressure; consequently, it is concluded that the flow through the exhaust system was unaf'f'ected by engine operating condition or Reynolds number. .

Effect of Compressor and Combustor on

Generalized Engi De Performance

The compressor and combustor were shown to be the only components. which had sign1f'icant Ilerforma.nce variations over a wide range of alti­tudes. The manner in "Which these component.Ilerformance variations are reflected in the over-all engine Ilerformance are next considered.

The shift in compressor operating Iloint with altitude for a given compressor Mach number is developed in figures 18 and 19 for an axial­and a centrifugal-flow compressor' (engines B and F, respectively). The shift in compressor corrected speed line with altitude, which was pre-

. viously discussed, is shown in figures 18(a) and 19 (a) .

A compressor. being run in an engine will operat~ at only one Ilres­sure ratio and air flow for arry given sIleed and altitude; consequently, as the engine BI>eed is varied the compressor operates along an ollerating

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line which passes through the compressor-characteristic map. With an increase in altitude; the operating lines for- both compressors shif'ted to higher pressure ratios _at a given corrected air flow (figs. 18(b) and (19(b». This upward shift in the operating lines resUlts from the reduction in com;pre-ssor efficiency at al ti tude; more power from the tur­bine per pound of air flow is then required to drive the compressor. The interaction of the compressor, turbine, and fixed-area exhaust nozzle . resulting from this increased demand on the turbine requires that the com­pressor pressure ratio at any given corrected air flow be increased to maintain an equilibrium operating condition.

Superimposing the corrected compressor speed and operat:!-ng lines for each com;presso:r;' (figs. 18(c) and 19(~) >. gave the operating points at sea level. A and at an altitude_of 50., 000 .feet B.Tiieaxia1-flow com­pressor operated along the steep portion of-the compressor corrected speed line. Consequently the. upward shift of the operating line combined with the shift in the corrected speed li~ as the altitude was increased moved the compressor operating point to a higher pressure ratio and lower corrected air flow • The centrifugal-flciw- ·c6iopressor , on- the other hand, operated on the more nearly horizant81portion-61 the compressor corrected speed line. Because of this characteristic., the downward shift in the corrected speed line with increased altitude more than offset the upward shift in operating line with the result that both the air flow and pres­sure ratio at the operating point were reduced by an increase in altitude. These shifts in compressor operating-p01n~-withan·accampanYing reduction in compressor effici-ency at high altitude are reflected in the over-all engine performance.

The effects of these changes in compressor operating point, with attendant changes in· turbine operating point, and-the reductions in com­bustion efficiency at altitude on over-all performance of both an axial­and a centrifugal-flow engine aJ;'e illustrated in figures 20 to 23. These figures present the primary performance variables of engines B and. F as functions of Reynolds nulnber index for constant corrected engine speeds and flight Mach numbers. Tbeperformance variables include the corrected jet thrust parameter, the correc~d aJ-r :f;lowJ the _ corrected turbine­outlet total temperature, and corrected fuel flow. Each variable is presented as the ratio of the corrected &ltitude value to the value at sea level.

As altitude was increased and the Reynolds number index reduced, the corrected jet thrust parameter was increased for the axial-flow engine and, except at one condition, -was reduced for the centr:L:f'"ugal-flow engine (fig. 20), although the corrected air· flow (fig. 21) was reduced for both engines. The shift in opere,tipg point of the axial-flow compressor to a higher pressure ratio assha-wn: in figure 18(c) and the reduction in effi­ciency were reflected in an' increase' in turbin~-ou:het total temperature at reduced Reynolds number indices (figs. 22(a) to 22(c», which was

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accomp?-Died by an increase in turbine-outlet total pressure. These increases with altitude more than offset the reduced air flow with the resulting rise in corrected jet thrust parameter. The shift in opera­ting point of the centrifugal-flow compressor to a lower pressure ratio (fig. 19(c», in ad.d.:1.tion to the reduced efficiency (fig. 1.2).1 resulted in an increase in corrected turbine-outlet temperature only at very low Reynolds number indices for the two lower engine ·speeds (figs. 22(e) and 22(1'» and a decrease in turbine-outlet total pressure at each speed. The combined reduction in corrected air flow and turbine-outlet total. pressure offset the rise in turbine-outlet temperature, which thereby· accounts for the drop in corrected jet thrust parameter of this engine at high altitudes.

The decreases in both compressor and combustion efficiency resulted in the increase in corrected fuel flow at low Reynolds number indices for both engines (fig. 23). The majority of this increase in corrected fuel. flow resulted from reduced combustion efficiency. Analysis of the data indicated that adjusting the fuel flows to unity combustion effi­ciency for all conditions greatly reduced the severity of the correct€!d fuel flow variation nth ReYnoldS number index. In one case, rated speed with the centrifugal-flow engine, the corrected fuel flow adjusted to unity combustion efficiency actually decreased with decreasing Reynolds number index. This reduction resulted from. the reduced corrected air flow and constant turbine-outlet temperature for that operating condition.

The variation of generalized performance with altitude or Reynolds number index for other current axial- and centrifugal-fiow turbojet engines bas been shown elsewhere (references 2, 3, and 15 to 18) to exhibit trends similar to those shown for engines :a and F. Therefore, although the ma:gni tude of the varia tiona differ from. one engine to another, and in one case between two engines of the same design (refer­ence 3), the trends illustrated here are typical.

Comparison of Corrected Engine Speed and Reynolds Number

Effects on Altitude Performance

An effect of altitude which is generally overlooked in discussing performance of isolated turbojet engines is the diffe+Snce in performance variations, primarily, corrected thrust, with corrected engine speed from one engine to another. Such differences in the slope of the corrected

~ net thrust curve plotted against corrected engine sp.eed is· often of such magnitude to completely overshadow the Reynolds number effects.

These differences are illustrated by the variation of corrected net thrust with· corrected engine speed, shown in figure 24 for centrif'ugal­now engine F and axial-flow engine B, the two engines for which perfor.m.­a:nce data were presented in :figtllcs 5 .to 8. The corrected tb.rUst is

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presented as a fraction of sea-level thrust at rated speed and the cor­rected speed is presented as a fraction of rated speed. Consequently, the sea-level thrust ratio for both engines is unity a.t rated corrected engine speed. The Corrected thrusts at sea level and at an altitude of 50,000 feet are shown for -engiIies F 8.ncl .B at- i'll-Slit Mach numbers of 0.67 and 0.62, respectively. This small dif'ferenc~ iIJ. flight Mach number has an insignificant effect on the sloPe of the curves. As altitude is increased 'With an attendant reduction in ambient air temperature, the corrected speed and corrected thrust increase, until at an altitude of 50,000 feet the corrected speed ,ratio is 1.14 when the engines are oper­ating at rated speed. A p~edicted carreC'ted thr'1:tst at ra.ted speed and based on sea-level a.8.ta is" indicated at -thls--operating point on the sea­level curve. Because of the Reynolds number effects previously discuseed, the thrust curve for an altitude of 50,000 feet i~ displaced as shown and the "actual thrust available occurred at slightly less than the rated speed condition for both engines, because of the turbine-temperature limitations.

It is apparent from. these data that at an altitude of 50,000 feet both the actual and predicted corrected thrust of the centrifugal-flow engine reached a considerably higher percent of sea-leve~ thrust than for the axial-flow engine. In the succeeding discussion, the magnitude of these differences in the ratio of corrected altitude thrust to sea­level thrust will be compared 'With the m.e.gni tude of the Reynolds number effect on maximum thrust at altitude.

Reynolds number effects. ~ The reduction in engine speed required as altitude was increased to avoid excessive- tu.rbine temperatures is -shown in figure 25(a) for five engines (A, B, C, G, and F). In each case the temperature-limited speed vas restricted below the rated speed as altitude was increased; the reduction was 1 to 10 percent for the five engines at an altitude of 45,000 feet. The attendant variation in maxi­mum thrust with altitude compared with the value predicted fram sea-level data is shown in figure 25(b). For each engine, except engine B, the maximum thrust fell-below the predicted value and, at an altitude of 45,000 feet, varied from 1/2 percent above to aa much as 14 percent below the predicted value. The opJ?osite trend for engine B resulted fram the upward shift in the corrected thrust curve with increased altitude shown in figure 24(b), which more than compensated-for the effect of the reduc­tion in :ma.xim:um speed. -Differences inmagni tude of the trends among engines were dependent on the amount the air flow was af'fected by speed reductions and the manner in which the - compress-or operating point was shifted by the Reynolds number effects.

Variation of the ratio of act~ to predicted specific fuel con­sumption -with altitude is shown in figure 25(c) for the sa:me conditions. Because of reductions in both compressor efficiency and coinbustion effi­ciency, the specific fuel consumption of each engine,_except engine A, increased at high altitudes above the predicted·value. With this excep­tion, the specific fuel consumption at an altitude of 45,000 _ f'eet was

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NA.CA EM E5lJl5 • 5

3 to 20 percent higher than the predicted values. The reduction ~ spe­cific fuel consumption for ~ne A resulted primarily from the fact that the cOIlJflressor operated at a considerably higher efficiency at the reduced speed at altitude than that predicted from sea-leVel data for rated speed operation.

It might be expected that use of a variable-area exhaust nozzle, which permits operation at rated speed and limiting turbine teIlJflerature at all altitudes, would reduce or'eliminate this difference between actual and predicted thrust ,at altitude. In order to e:x;plore this pos­sibility, variation w.l..th altitude of the ratio of maximum thrust with a variable-area nozzle to ma.x:imum obtainable thrust with a fixed-area noz­zle is shown for the five engines ~ figure 26(a). Exhaust-nozzle vel­ocity coefficients of the fixed- and variable-area nozzles were assumed equal. These data indicate little or no gain in ma.x:imum tbrust at alti­tude by use of the variable-area nozzle, except for engine C, which ga.~ed 4 percent in thrust at an altitude of 45,000 feet. Analysis of the performance of this engine indicated that at a higher flight Mach number, cOIlJflarable with that of the other engines, most of this gain would vanish. Comparison of the specif'ic fuel consUIllptions in figure 26 (b) shows that for all but one engine the specif'ic fuel consUIllption at high al ti tudes was higher for rated speed operation at limiting temperature than at the teIIJflerature-limited speed with a fixed nozzle •

These trends can be e:x;plained by use of figure 27 which presents the variation of corrected air flow and compressor efficiency nth corrected engine speed for each engine at high altitude. The teIIJflerature-limited speed of each engine is indicated on the figure by ticks on the' curves and the common value of rated speed is ~dicated by the vertical ~e. Although the air flow for each eng~e rose slightly with the ~crease from temperature-limited speed to rated speed (fig. 25(a), the compres­sor efficiency was reduced because the e~es were operating on the negative slope of the compressor efficiency curves. The consequent reduction in compressor efficiency for operation in this re&1on resulted in a decrease in total pressure ratio across the engine P4fPr In most cases, this decrease in engine pressure ratio nearly offset the increase in air flow thereby resulting in little or no increase in thrust and a rise in specific fuel cons~tion (fig. 26).

'When the added weight, cOIlJfl li cation, and possible lower velocity coefficients of the variable-area nozzle as compared with those of a fixed-area nozzle are considered, there is no apparent advantage in using a variable-area nozzle to permit operation at rated engine speed at alti­tude. ActUally, the variable-area nozzle wouid be detrimental in some cases, because of the rise in spectl.f'ic fuel consumption.

Corrected engine speed effect-so - Consideration will next be given to the effect of corrected engine speed on the corrected performance at al ti tude and the magnitude of differences in this effect from one engine

14 ••. -- NACA RM ESl.n5

to another as compared with the Reynolds number effects. In order to isolate the corrected speed effect, the variation with altitude of the ratio of maximum corrected net thruBt predicted at altitude to sea-level thrust at rated ~eed is shown in figure 28(a). These data illustrate the increase uf this corrected netthrtist_ rat~Q ~t~_alt~tude as corre~ted engine speed was increased-by the reduction in ambient air temperature. As shown, the thrust ratio for centrif'ugal-flow engines F and G increased to slightly over 1.70 above the _ tropopause as cOD'I.Qared "With approximately 1.35 for axial-flow engines A and B. In: other words, on the basis of . 'sea-level performance data, if the sea-level_tnrusts of t~e~e engines were equal, thrust of ,engines F and G above the trOIlopause would be approxi­mately 26 percent _hlgherthan that of engines A and. B. The data for engine C are not directly comparable- with the other data because of the difference in flight Mach number. Examination of the performance data indicated that raising the flight Mach number :from, 0.25 to 0.60 WOUld reduce the thrust ratio of _.tPis engine ~boye the t~oJ?opause by approxi-mately 5 percent because of dif':('eren~es in sl~e of the--net thrust ---curves.

For each engine the ratio o~predicted to sea-level specific fuel consum;ptions decreased up to the tropopause (fig. 28(b», except for engine A, which suffered a severe reduction in compressor efficiency at high corrected engine speeds. In general, the reduction in specific fuel consumption was greater for the engines having the greatest rise in pr-e-­dicted to sea-level thrust ratio with an_ increase ;I.n altitude~

Combining the ,Reynolds number effect shown in figure 25 (b) with the corrected speed effect (fig. 28(a», to give_the ratio Cl:f'corrected maxi­mum net thrust at altitude to sea-level net thrust at rated speed, results in the trends shown in figure 29(a). At an altitude of' 45,000 feet the thrust ratios for the axial-flaw engines fe],l betw!3en 1. 23 to 1.34 as compared with 1.53 to 1.61 for thecentrifug8J.-flow engines. - Corres­pondingly, if all engines had _ the same .thrust at sea level, the centrifugal-flow engines would give 14 to 31 percent more thrust at an altitude of 45,000 feet than the axial-flow eng1D.es.

The ratio of actual to sea~level specific fuel consumption remained essentially constant up to an altitude_of about- 30,000 feet for the axial­flow engines and decreased rather rapidly_for the centrifugal-flow engines (fig. 29 (b ) ) • Above this altitude'; thesPeci:fic fuel' consumPtion increased very markedly for the centrifugal-flow engines and more gra.d lla 1 ] Y for the axial-flow engines. _ This.in~ease _ was mainly due to reduced compressor and combustor efficiencies. - '-.,---... - ,-'"----. .. -- -- - --

The difi'erences in the trends of tbe thrust rati9 between the axial­and centrifugal-flow engines is not inherent in the respective compressor tY.J;les, but is _aSSOCiated -with the relative air flow per unit flow area of' the cOIDpressors when they are operat~ at rated corrected speed and above.

•

•

N (.-~

-N en

-~

--.:......::,.j..

JI!. _

NACA EM ESlJl5 l5

This comparison is illustrated by the corrected air now and compressor efficiency shown for each of the five engines in I'igure 27. These data show that the corrected air flow for the centr1:fu.gaJ.-flow engines con­tinued to increase more ra:pidly than that for the a.xia.l-,flow engines above rated corrected ~eed~ which accounts for a small :portion of the di:f.'ference in the ratio of: altitude to sea-level thrust. The difference in air flow characteristics results from the fact that the axial-flow com;pressors, :parti,cularly "tflose of engines A and B, had a very high air flow :per unit area and were J:!.ear choking in the early stages at rated ~eed, whereas the compressors of eng:tnes F and G were still operating below the choked conditiOn at this s:peed.

The manner in which differences in com;pressor characteristics among the engines, and associated. interactions of the other components a:;E'i'ect the total :pressure ratio P4!Pl and total ~erature ratio T4/Tl

across these engines is shown in figure 30. As illustrated, the engine :pressure ratios rose considerably more rapidly above rated corrected speed for the centrif'ugaJ.-f'low engines than for the axial-flow engines. Although there was no consistent trend in engine temperature ratio with engine type, the effect of this variable on thrust is considerably less than that of engine pressure ratio. It is the summation of these dif­ferences in trends from one engine to another, in addition to dif'f'erences in air flow characteristics, which account for the markedd1f'ferences in the ratio of maximum altitude corrected thrust to rated sea-level thrust.

It may, therefore be e~ected that an engine designed for a high air flow :per unit flow area, and consequently operating with the flow restric­ted by inlet choking at rated corrected speed and above, will have a con­siderably lower thrust' at high altitude than an engine with equal sea­level thrust, but designed for a sufficiently low air flow :per un1 t flow area to :permit operation at rated corrected ~eed or above with the iillet unchoked. This characteristic is an important consideration both in engine design and :in selection of an engine for a. high-altitude aircraft.

CONCLUDING REMARKS

The compressor and the combustor are the only two engine components found to have performance characteristics which were affected over a wide range of altitudes. Two engines exhibited a slight decrease in turbine efficiency, and two had a slight rise in combustor total-pressure-loss ratio at very low engine-inlet pressures~ corresponding to flight at alti­tudes above 50,000 feet. The ~duction :in compressor efficiency and cor­rected air flow with Reynolds number index and' the reduction in combustion . efficiency with reduced altitude pressure are the factors which result in failure of' engine_ p~rf'orma.nce 'to generaJ.ize f'or all altitudes and which therefore have a detrimental effect on the altitude performance of turbo­jet engines. In general, the effect is to increase the turbine-outlet

5

16 ' ,0 NACA RM E5lJl5

temperature at rated engine speed as the altitude is raised; a speed reduction is thus required to avoid exceeding 'the turbine temperature limit. The' attendant r~ductions in c~ressor e£ficiency and corrected air £lowat reduced Reynolds number indices reduce the thrust below the value predicted £rom a. sea-level performance calibration. Also, the combined e£fects of reduced cOmpressor and combustor efficiencies at high al ti tudes result in an a];lIlreciable increase in speci£ic fuel con­sumption above the predicted values.

An al ti tude e££ect which was f'o~ to be of considerably greater importance than the effect of Reynolds nu.niber index 'was" the corrected engine s];leed ef£ect. It was found that the rate of increase in corrected net thrust with corrected engine speed, and consequent~ the rate of thrust decay withal ti tude -' differed considerably from one engine to another. When the five engines discussed were assumed to have equal thrusts at sea level; the ma.xiriiuin. 'thrust at an altitude of 45,000 feet d1£f'ered f'rom one engine to another by as much as 26 :percent because of the corrected engine' speed e£fect alone. The combined corrected engine speed effect and Re~olds number ef'iects resulted in differences in max­imum. thrust at this altitude of as much 'as 31 percent between the engines having the highest and lowest thrusts. The corrected engine s:peed effect is attributable to the compressor air flow and ef£iciency characteristics and attendant interaction of the other com;ponents. The engines with com­pressors designed to handle a higlJ. air flow per unit flow area, and 'Which consequently have some of the stages choked at hfgb. engine speeds, exhib­ited a more rapid thrust decay with altitude than those designed for lower air flows. It is therei'ore' ~ortant that these characteristics be considered when the initial design Criteria for an engine and. com­pressor are being established. Furthermore, when an engine is selected for a specific mission, the 'choice should be based on the performance at design altitude rather than on the sea-level rated thrust and spec1£ic fuel consumption.

Lewi s Flight Pro~sion Laboratory National Advisory Camm1tiee for Aeronautics

Cleveland, Ohio.

l'

..

a,.---

NACA EM E51Jl5 17

APPENDIX -CALCUIATIONS

Symbols

The f'ollowing symbols are used in this reIJort:

A cross-sectional area, sq f't

B thrust scale or thrUst diaphragm. reading, lb

D external drag of' installation, lb

d diameter, f't

F j jet thrust, lb

Fn net thrtis~,- 1b

f'/a fuel-air ratio

g acceleration due to gravity, 32.2 f't/sec2

·R total enthalpy, Btu/lb

R:r I enthalpy of' fuel components in product!;! of' cc:anbustion, Btu/lb

hc lower heating value of' fuel, Btu! lb

M Mach number

N engine speed, rpm

P total pressure, lb/sq f't absolute

p static pressure, lb/sq f't absolute

R universal gas- constant; 53.4 f't-lb/(lb)(~)

Re Reynolds number

T total temperature, ~

Ti indicated temperature, ~.

t static temperature, "R-

V velocity, f't/sec

18 3 2 NACA RM E5J.Jl5

air flow, lb/sec

Wf fuel flow, lb/b:r

Wg gas flow, lb/sec

y ratio of ~ecific heats for gases

8 ratio of engine inlet absolute total pressure to absolute total pressure of NACA standard. atni.os:pher~ at sea level

T} efficiency

e ratio of engine inlet absolute total temperature to absolute total t~erature of NACA standard. atmosphere at sea level

~ absolute viscosity, slugs/ft sec

p density, slugs/cu ft

~ ratio of air absolute viscosity at- engine-inlet total t~erature to air absolute viscosity at total temperature of NACA standard atmosphere at sea level

SubscriRts:

0 free stream

1 engine inlet

2 compressor outlet

3 turbine inlet

4 turbine outlet

5 tail-pipe nozzle outlet

a air

b combustor

c compressor

e engine

f fuel

--------'----

..

NACA RM E51Jl5

m fuel manifold .

s tail-pipe seal at bulkhe~

t tUJ:'bine

:x: inlet duct at slip joint

Generalized Performance Variables

The generalized engine perf'ormance variables used in this report are defined as follows:

Ji'n/6

N/.,fiJ

T4/e

Wa -.;e 6

Wf /(6 -/8)

Wf/Fn...Je

corrected jet thrust parameter, Ib

corrected net thrust, Ib

.-corrected engine speed, rpm

corrected turbine-9utlet total temperature, OR

corrected air flow, lb/sec

corrected fuel flow, lb/hr

corrected net thrust,specific fuel consumption, Ib/(hr)(lb thrust)

Methods of Calculation

19

Total temperature. - Total temperatures were calculated from the indicated temperatures, using a thermocouple recovery factor of 0.85, and the values of' pressure· and ratio of' specific heats at the respective stations in the engines.

T -7-1

. 7 1 + 0.85 (~) - 1

(1)

20 • NACA 1M E5lJl5

Air flow. - For the axial-flow engines, the air flow was calculated. from p;t'essure and temperature measurements obta1.D.ed at the engine inlet (station 1) ,

(2)

The same expression waa used. for calculating the centrifugal flow engine air flows with the tail..,..pipe outlet measurements (station 5) substituted. in place of the engine inlet values.

Compressor efficiency. - The adiabatic temperature-rise compressor efficiency f'or both the full-s®le, engine a.nA, compressor component inBtal­latio:o.a was calculated by, the following equai;:.ion:'

-(~?f!--~.-'Ilc = T2

- -1 Tl

(4)

where r c is the ratio of specific heats corresponding to the average total temperature of the ,air flowing through th~ C?om.press or.

Turbine efficiency. - Turbine efficiency was calculated from. the equation:

rat = (5)

'Where 7t is the ratio of specific heats corresponding to the average total temperature a:r the gaa flowing throu~ the turbine. To obtain turbine-inlet temperature, the enthalpy at the turbine inlet was calcu­lated from the turbine-outlet temperat'ure and compressor temperature rise; campr~saorand turbine work were assuIDed equal.

Turbine-inlet temperature was then determined by the use of enthalpy charts.

(6)

~. --.

-- ...........

.. i ..

RACA RM. E5lJ15 21

Combustion efficiency. - Combustion efficiency was calculated from the enthalpy rise across the combustors by the following equation:

Ha,3 + f/a (Hl f ,3 - Bf ,m) - B8~2 ~b = fJa hc (7)

The enthalpy of the fuel components in the products of combustion H lf 3 ,

vas d.etennined .from the hyd.rogen-carbon ratio of the f'uel and. the calcu-lated turbine-inlet temperature by the method, d.escribed. in reference 19.

Thrust. - For engines installed. in the altitude wind. tunnel, the measured. thrust was 'd.eterm1ned. from balance scale measurements by use of' the relation -

Wa F j = B + D + g V:x: + Ax (Px -- PO)

where the externald.rag D was d.etermined. .from. power-of'f' I

Because there was no measurable external d.rag or initial momentum f'or engines installed. in the altitud.e chamber, the measured jet thrust was d.eter.mined. f'rom force measurements on the thrust d.iaphragm by the relation -

(9)

where PI 'W'8.8 taken as the static pr~sure in the inlet section of the chamber.

The calculated. jet thrust, which was d.etermined. f'rom pressure and. temperature measurements at the turbine outlet, was calculated by the relation

274, . [ (PO)7~:1 '-7.-4---1- ~4, 1 - P-4,- (10)

Net thrust for eaoh engine was calculated. using the measured. thrust in the follow1.ng equation:

(11)

-where f'ree-stream velocity Vo was calculated from engine inlet total

pressure PI and. temperat~ Tl and. f'ree-stream statio pressure PO~ . assuming oomplete free-streB.1ll ram pressure recovery at the engine inlet.

22 NACA EM E5lJl5

Reynolds number -index. - In developing the expression for Reynolda number index (reference 8), Reynolds number of the engine may be express ed. as:

Re - (12)

In the case of' a turbojet engine, it is convenient to consider the mass velocity PIVI as the mass flo}i per unit frontal area Wa/A.a. The

Reynolds number is then

(13)

Reynolds number is not a convenient parameter because mass flow is a a function not only of the pressure and temperature of" the inlet air, but is aleo dependent on two other independent variable~, corrected engine speed and. compressor pressure ratio. A more convenient parameter for considering the affects of fluid viscosity is

(14)

This parameter is related to the ratio of' Reynolds number ;to Mach" number and its value depends only on the condition of' the air at the engine inlet. In generalizing et;lSine data, itie convenient to use 8 in place of' P and 6 in place of' T. When the same reasoning" that led to the development of' 8 and 6 is f'oilowed., another quantity q>, is defined as : ~ = viscosity at engine-inlet ~otal temperature

viscosity at total temperature of NACA standard sea-level atmosphere (is)

When the quantities 8, 6, and <P are used in equation (14) in place of P, T, and" IJ., respectively, and the constants de, y, and R are omitted, the followi,ng form. of' Reynolds number index is obt~ined:

Reynolds number index = r;:;:-IP,,6

8 (16)

The value of this Reynolds number index is 1.0 at NACA standard sea-level pressure and temperature.

RHlli'ERENCE3

1. Sanders, Newell D.: Perfonnance Pa,r.ametere" for Jet-Propulsion Engines. NACA TN 1106, 1946.

..

~ --

. --

.-~ .........

_ T

N ()I N en

NACA EM ESlJlS 6 25

2. Vincent, K. R., and Gal.e, B. M.: Altitude Perf'ormance of' J3S-A-11 Turbo jet Engine in Al ti tude Chamber. NACA EM ESOnS, 19S1.

3. Wi1sted, R. D., and Grey, R. E.: Altitude-Perf'ormance and Reynolds Number Investigation of' Centri~l-Flow-Compressor Turbojet Engine. mCA RM E51B01, 19S1.

4. Wallner, Lewis E., and Fleming, William. A.: Reynolds Number Eff'ect on Axial-Flow Compressor Perf'o:r:ma.nce.NACA EM ESGll, 1949.

S. Sanders, Newell D. and Behun, Michael: Generalization of' Turbojet­Engine Perf'ormance in Terms of' Pumping Characteristics. NACA TN 1921, 1949.

6. Prince, William. R., and Jansen, Emmert T.: Altitude-Wind-Tunnel Investigation of Compressor Perf'ormance on J41 Turbojet Engine. NACA RM E9G28 , 1949.

7. Dietz, Robert 0., Berdysz, Joseph J., and Roward, Ephraim M.: Altitude-Wind-Tunnel Investigation of' a 3OOO-Pound Thrust Axial­Flow Turbojet Engine. II - Analysis of' Compressor Perf'ormance. NACA RM EBA26a, 1948.

8. Dietz, Robert 0., and Suozzi, Frank L. : Altitude-Wind-Tunnel Investigation of' a 4000-Pound~Thrust Axial-Flow Turbojet Engine. IV - .Analysis of' Compressor Perf'ormance. NACA EM EBF09c, 1948 •.

9. Dietz, Robert 0., Jr., and Geisenheyner, Robert M.: Investi~ation

of' the I-40 Jet-Propulsion Engine in the Cleveland Altitude Wind Tunnel. II - Analysis of' Compressor Perf'ormance Characteristics. NACA EM ESG02a, 1948.

10. Thorman, R. Carl, and McAulay, John E.: Altitude-Wind-Tunnel Investigation of' Turbine Performance in J47 Turbojet ~ne. NACA EM E9IG.0, 1950.

11. Krebs, Richard P., and Hensley, Reece V.: Altitude-Wind-Tunnel Investigation of' a 4000-Found-Thrust AXial-Flow Turbojet Engine. V - Analysis of' Turbine Performance. NACA EM E8F09d., 1948.

12. Campbell, Carl E.: Altitude-Wind-Tunnel Investigation on Combustion Chamber Perf'ormance J41 Turbojet Engine. NACA EM E9L02, 19S0.

13. Childs, J. Roward: Preliminary Correlation of' Efficiency of' Aircraft Gas-Turbine Combustors for Dif'ferent Operating Conditions. NilCA RM E50FlS, 1950.

24 NACA RM ESlJlS

14.· Campbell, Ca.;!:':L.:E.; A:L:ti tud,~-Rin,c:l~~J"J:nv~s.tigl:l.t:!:9~ of,' 3009.-Po~d-. Thrust Axial~Flow Turbojet Engine. _ III - Analysis of Combustion­Chamber Performance. NACA RM EBB19, 1948.

lS~ -- Chnxad, E. William, and Sobolewski, Adam E.: Altitude-Wind-Tunnel Investigation -.of J47 Turbojet-Engine' Performan~e. _ NACA RM E9G09, 1949. --

16. Meyer, Carl L., and Bloomer, Harry E.: Alti tude-Wi nd-Tunne1 Investigation of Performance and Windmilling Drag Characteristics of' Westinghouse X24C-4B Axial-Flow Turbojet Engine. NACA RM EBJ2Sa, 1948:

17. Bar SOh; Zelmar, and Wilsted, A. D.: Al ti tude-Chamber Performance of' British Rolls-Royce Nene II Engine. I - Standard lCl.7S-Inch­Diameter Jet Nozzle. NACA RM E9I23, 1949.

18. Gendler, stanley L., and Kof'f'el, William K.: Investigation of the 1-40 Jet Propulsion Engine in the Cleveland Altitude Wind Tunnel. I - Performance and Windmilling Drag Characteristics. NACA RM E8G02, 1948 •. --- --- ------- --- ---

19. Turner, L. Richard, and Bogart, Donald: Constant-Pressure Combustion Charts Including Effects of' Diluent -Ad<lition. NACA Rep. 937, 1949. (Formerly NACA TN IS 1086 _and 1655.)

--.. ----~

~ 1. ~ :I:nB"tAll.at1on at' tlzrbo.jet ~ bJ. alt1tl!db lI1nd. t=l.

~ . , .

, .

I':'~

~t t l

.' O?~ .'

..

:Fig\u'e 5. - Setup fo:r: 1nTest1aat1ng perf()l"!N1OA at =t&l-flow oanpreSIKll'.

28 NACA RM E5lJ15

Air flow

(a) Statio pressure tube instrumentation at oompressor inlet.

Thermocouple

Total-pressure tube

Air flow

(b) Total-pressure tube and thermocouple instrumentation at oompressor inlet.

Figure 4. - Typical survey r~e installations.

NAeA RM E5lJl5

Trailing edge or compressor-outlet straightening vane

(c) Static -pressure tube iDStrumentetion at compressor outlet.

Thennocouple -'--"'---

compreasor­outlet straightenillg vanes

Air f'low

(d) Total-pressure tube and thermocQuple inBtrumentation at compressor -outlet.

Figure 4. - Cont1nued. Typ1cal rake installations.

&

29

30 NACA EM E5J.Jl5

Air .flow

(e) Total-pressure tube instrumentation at turbine inlet.

Figure 4. .,. Continued.. Typioal survey rake installations.

92£2 , .

'?

~ J1111l1111 11 . " "'

•

l ThermoooUPle

. Air flOIf'

.. A

Total-pressure tube ~

1-.. l

-c:- -.............

~

-(t) T~-preB8ure tube ani thmnoooup1e :l.nBtl'ImIentat1on at. tul-b:lne outl"t.

:B':!.gure 4.. - Continued. T1P1olll. B1U'VeJ" rll.le :!Ilatal].at1oDB.

·32

-Exhaust nozzl wall

e

Air flow

Thermocouple

..

stati c-pressure tube.

Total -pressure tube

NACA RM E5lJl5

.~ r-II

~.

I =- =.:z::.-1/

_ .

-

~ -~

/

~.

(g) Total- and static-pressure tubes apd thermocouple instrumentation at exhaust-nozzle outlet.

Figure 4. - Concluded. Typical survey rake installations.

~ " .. '" ...... . ~_:..:.-~.3

5Q.

}

J ') J

NACA EM E51Jl5

l80r---~---4----+----+----~--~

160r-- 0

Altitude (f't)

10,000 o 20,000

1------1 <> 30,000 r----+---+---i " 40,000 6. 50,000 f

l'Or-~_r--_.--~+_--_+----~~~

I( i fJ 'i-4 120 I-----+-----f----+----+-.l-I'ci-F--"+-1'-__ ----1

i$ ~ f-~ 100

j ~ ~ 801-----+-----f----+-~/+-+----+-----+

I a o .~

601-----+-----f----~~---+----+----~

/

j-40r---_r---4~-4~+----+----+---~

J ...... 20~ __ -L __ O~-~ __ ~ __ ~~ __ L-__ ~

60 80 100 120

~ IL

II! j

60 'V 80 Corrected engine speed, :percerrt of rated

J r

~ ~

>

~ # ~

W

~ 100 120

(a.) Engine B, axial-now engine; (b) Engine F, centrif'ugal-i'lDw engine; fltSht Mach number, 0.62. flLght Mach number, 0.67.

Figure 5. - Ef"i'e.ct of' altitude on corrected net thrust.

33

34

il 120

~ 'd ~ 8 100 f.<

~ ~

~ ~ 80 ~

"Cl $ 0 II)

~ t.>

60

2.4

~f 2.2

~

2.0 .; 0 ..... ..., ~ ~ 0 1.8 tJ

.-!

~ tJ ..... .... ...

1.6 tJ t> ~

~ t t>

~ 0 1.4 t.>

1.2

NACA EM ES1J15

10-0 0 0 "l A

br f.>"

.A

? V .tt )

V l:l /

~

(~ 80 _lOo uo GO eo - ~po -120

Corrected eoe:1ne speed, percent of rated

(a.) Eosine B, -nial-rlow ens1nB; (b) EIJgIne 7, centr11'uaal-flow eoe:Ine; fl1sht Mach number J 0.62.- fl1ght M!l.ch nUlllber, 0.67.

Figure 6. - Effect of altitude on corrected a.Ir flow.

lltltude (tt)

0 10,000 I---

t 4 0 20,000

_. 0 __ ~.o,OOQ t----"l 40,000 A 50,000

1\

\ r\ \ '\

\ ~ h

0 \

\ \ ~'\7 \~ ,\\ ~\

~

h\ ~ p ~ ~ ~ \ ~

J. V M b ~

, ~I 60 80 120 6Q - - - - 80 ).20

Corrected engine speed, percent of rated

(a.) Engine B, ex1al-fl.owengine) (b) Engine F, centrlfu&al-tlow engine) flight Mach number, 0-.62. :t:light Mach number, 0.67.

Figure 7. - Effect of .altitude on corrected ~pecifl~ tuel cons,ump~ion.

- --~­~

:10" - -

NACA BM E5lJl5

2400 1-----1

A1titude (ft)

o 10,000

OD 20,000

\--->-...; :30,000 -1-----1----+----..1 "V 40,000 L:!.. 50,000 l::::.

2200 I-----l-------.>----_+----+----+-Ih~'--I

cF.. fV ~ 2000 ~/I

J 1Tf!' j 1800 zVJ .p

~ ] i 1600 711 : .~

.8 1400 1----"-' +--_-+-_-wrJ_--+_--f_---+

I ~ 1200 I----I-----I~"____=_J J1'____.+_--_+----+__--__f

Cl u

35

-

-

f / ~

1 >

1/ f ,r

It! (,j

~ !P

..d V I

V .;J

~ ~

1oooL-__ -L __ ~~~J---~----~--~ 60 - 80 100 120 60 80 100· 120

Corrected engine speed, percent of: rated

(a) E:ogine B, ax1a1.-f:1ow engina; (b) E:og1ne F, centrifuSal--f:1ow engine; f:11ght Mach number, 0.62 night Mach number, 0.67.

Figure B. - Effect of: altitude on correoted turbine-out1et total. temperature.

0 .... ~ ,., ID

~ ., ., oil

!l. a .. ., CD

f <:>

3.4 r_--1----+~~~--~--~--_+----r_--4_--_+--_1----+_--1_--_+--~~--+_--~--~ Com.preesar-iDlet

3.0

2.6

v: 2.2

1.8

1.4 65

75·

total :pr9sBllre (lb/sq tt)

-0- au --0-- : 422

n-nh~ []

70

\ 80

15 80

89

85 90 Corrected sir tlOll, llE'rcent of rated

(11) Engine C, a.:daJ.-flow oampressor.

95 100

Figure 9 •. - lfi'fect of 1lJlet pressure on cbara,cter1st1os of compressors 1l!Veot1ga ted in compresaor componeJIt U1sta.lJ.B,'t1one~

_~ I ..

I.

I

0 ..... .p

e ill

~ ID ID ill a. a III ID ill

t <:,)

4.5

:;.5

3.0'

2.5

2.0 ·h

1.6 30

UlI

0' [)75

Compressor-inlet total presam;oe

(lbjeq tt) -0-- 986

I --0-- 352

VI' I... / .............

K' / }---- .[J---I--~. ~ .. , / ~ ....

IUom.pressor -- // ~/ / V eff1oieDD~ ," -./ ",' /

.-// /' ~ ~ ,75 ;/' ,. // / " " ,

~7~ // ~ 7 / ~ / ,/ ,,"/

,I /' ~ // ~ /" 'j:, ,,-

'/~ ',/ 75 f ~ 7' ,/~ ~

( ,,- 85 ... '1l;Y' I)..... ~72? ..., ~D5

80 40 50 60 70 80

Correotl'ld air flow, percent of rl'!.ted

(b) Engine P, centrifugal-flow oompressor.

9&£2

-rr jt( 0.

h\ X ~ " ~ / ~

W I // / I( I

I

/ r/ /

100 Corrected com-pressor speed

percent of rated

~ 90 100

Figure 9. - Concltlded. Effeot of inlet pree81Il'6 on oharaoteristios of oanpreeeors imeBt1gl'!.ted 1n oompre BSor oauponent insta.l.la tions •

38 2 £ & NACA RM E5lJl5

85

8 .. ..... " .... ~~ () 8 80 ~ '" .. $. m ..

----p r! I --b--p/" V ~ Compressor-1nl.et. total

L tempera.ture

c:S r1 (<>R)

0 4,60 0 600

t 0

75 ! ! I (a.) Canpressor efficieno7.

H(lllifft Illob

[ R 200 4.00 600 BOO _ 1000

» 8 .. .... " .... ~~ '" '" " ~b .,p. ., tl r..

~ 0

~il ~~ -:;Ie,.., .-c 0

$1:l " ., II> !! ~ tl 8P.

Compressor-inlet t9ia.J. pressure, 1b/8Il. :rt

(b) Correoted a.ir.flow.

Figure 10. --Variation of axial-flaw-compressor performance with inlet. condit.ions. Correoted oompressor speed a.ni ocapressor pressure ratio, oonstant; engine A.

90 Corracted.~ressor speed

r 79 r 85 (percent o~ rated)

80

0 nl r2 LT ILQ) o·

~ .... /-T v ...C --

70

(a) Compressor efficiency.

100 Corr_eC-tedcompressor SJ.)l'led

(percent of ratl'ld). Q? ... ,.. b u

90 85 I,....,

~ .... - [J

Inlet. total Inl.et total

eo 19 ~o

temperature pressure (OR) (llJjsq ft)

d' ~ 0 4.60 5lZ 0 sao 1Z1

~ 70

:3 .0 :1.4 - 3.8 -. 4.2 - - --·'·.0 5.0 -Compressor pressure ratio_ ,-(b) Corrected air floy.

Figure U. - E«ect of inlet condit1.ons on u:!.al-fipv-'cjJmpresBor per:!'=. Reynolds num:;.'er- index, 0.282} ei1gI._ne A,. •

... _--

.-- ~

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

.,~":" :~

---7

. -- ............. J<, -.'

...... -

J J ") 1

NACA EM E5lJl5

cOmpressor Compressor type

A Axial. flow ------ C Axial flow ----- D Axial. now ----- E Centrif'ugal flow ----- F Centrifugal flow

100 --~ - ----I--:-? ~- ---- ~ ---V" --~

~> ~ ....

95

I 90

(a) Oampres~or efficiency.

100 ~ ~ ~- fo;-".:;o-

. /' ......... ~7 i'""" .......

.... .L ~ .. ~ ......

~V~ V~ , ....

I

~~ 90

a .2 .4 .6 .8 1.0

Reynolds number index, 19'\/8

(b) Corrected air flow.

Figure 12. - Variation of' axial- and centrif'ugal f:Low Compressor per­formance with Reynolds number index. Corrected compressor speed, rated; 'PI'essure ratio, constant. -

59

40

0 0 <>

.4.5

1'\ /'0 ~ ,-

'""' LlO

~ ~ f(

,; .35

/

i -.25

(a) Engine A, axial.-flow eng1ne)!lingle-stage turbine._

.50

III [ 11411 ,I 18 .«l

(b) ~ne B, axial-flaw engine) single-stage t~bine •

.50

M ,.A

" ~ V- .... ~

? 0

16 0 , - <>

.40

.30 (c) Engine D, axial.-fl.ow eng:I.nej single-stage turbine •

• 60

.... "'" ~ 0

--.g 0 -.;I 0 <>

.50

- -

~.4. , __ ,_, 1.8 2.2 2.6 3.0 Turbine pressure ratio

(d) Engine F, centrifugal-floYengine) single-stage turbine.-

I

I

NACA EM E5lJl5

Engine-inlet total. pressure '--

(~b/sq ft)

l.aU aoa -319

--rH -~~.-(lb/sq 1"t.) ~239 486 18

Erig!.ne-iruet total pressure -

(~b/sq ft)

-174.6 11.67 -

470

I I

Engine-inlet total. pressure r---(~b/sq :rt)

~625 678 r---262

I '-~

Figure ~. - Effect of al.ti tude pressure on corre~ted turbine gas_ f'lov.

---~

~.

- ...

-p

-r £ -'.~

.:.. .. _ .. :3.

6Q NACA EM ES1Jl5 • --...... - 41

90

:!ng1ne-1DJ..et total. pressure r--

(~b/sq 1't) 17h[ /'

~Bl.4. 1"'\ 0 r--- u v 0 8()8

[J 0 0 3l.9

80

70 -..., (a) E!lgm, A; ax1al.-fiow engine; siDgle-~ turbine.

90

EDs1ll&-~t total. pressure ~

(J.b/eq 1't)

~ 0 1239 I--0 71 t- O ~6

-' 0 J.Sl.

,... I I

EJJgm,-1DJ..et total. pressare t--

l.....o-c DQ-(n/sq1't)

~ ......" 0 J.74.6 t--

/ \'"' tj 0 1167

0 ~O

80

70 ~ I I (e) Ecg1ne D; a.z1al.-fiow engm,; siDgle-stse;e turb1ll&.

90

Ecg1De-,1nJ.et total. pressure

(n/ Stf. 1't) :l-. 0 l.S25 I--

~ jtO ~ 0 578 0- 0 262

80

<:f I ~ 'n 7 1.0 1..4. 1.8 2.2 2.6 3.0

TUrbine pressure rat10

(d) Engine 1'; centr1£usal.-1'l.ow eDg~; siDgle-stage tlIl"b1ne:

11.gUre U. - Effect of a.ltitooe pressure on ttn'b1Ile effioiency.

42 ---- NACA RM E5lJ15

100

leon ~ ~ I .... ....

90 I~ ~ p---~

i Corrected engine speed (perceDt of rated)

~ 0 100 f--0 90

80 0 80

(a) Engine Ai axiaL-flay engine J tubular combustor.

100 ...1"1

~ ;;t:-:;r

" '"

~ --- ~ ..."

~ ~ ( V V

.,/ [n

IV (b) Engine B, anaL-nov enginei tubular combustor.

, n. ..1"'1.

,/' V n ---10--~ r-o

V 7 ru -90

/ ~ H:> ~

[to <> ~ ac J .. . .. _ .

(c) EDg!.ne C, axiaL-flow eng:I.neJ annul.ar combustol:. 1

a,....-L--Q' L--......... 1-6

VJ V ~ ~

./ 90

/ ! V .?

Ii V

"" d ~ . ---- - -

o ~ 800 1200 1600 Engine-inlet total pressure, P

l, lb/sq t't

(d) Engine F I centr1:rugal-tlov engine J tubul.ar combustor.

Figure 15. - Variation at' combUliltioD. e:N'ic1ency nth engine-inlet total pressure at con­stant corrected engine speed :ror engines hartng tu"'aillar end BllIlUl.ar codlustors.

- .. --

NAeA RM E5lJl5 ----- 43

Corrected engine speed (percent of rated)

.06 0 100 0 90 <> 80

.A. "..

.04

~ ~ :v

.02

(a) Engine A. axial-flow engine; tubular comhustor.

] 11 ~ 1 r 1 I~ 1 1 1 1 1 1 (b) Engll:le B. axial-flow engine; tubular ccmhustor •

• 08 L'; .....

..It') ..., rv' ~ "V"

.... ....1""'1 n_ -u

... [

.06

'" ~ -cp.-o- -

.04

(c) Engine C. axial-flow engine; annuJ..a.r coinbustor.

:::I 1 1 EI:I I: 1 1 I: I~?I o 400 800 1200 1600 2000

Engine-inlet total pressure, PI

Cd) Engine F, centr1:f'uga.l-fiow engine; tubular combustor.

Figure 16. - Variation of combustor tOtal-pressure-loss rat10 with engine-inlet total presBure at constant corrected engine speeds for eng1nefl having tubular and annular combustors. .

44

1.0

.9

.8

1.0

.9

.8 1.0

<>~ ~ cY

V <> [J

)

0 nr

.P-~ ) 'V

1.4

NAeA EM E5lJ15

Free-stream static - .. -

0 CJ

<>

~ r

~ ~

-"!)

Ol.

(a) Eng1J:;te A, axial-flow eDgine.

,., A.

... U-I

C

<>

_1.8 2.2 2.6 Ttn"bine-outlet total pressure Free-stream static_pressure

(b) Engine C, axial-flow engine.

pressure (lb/sq ft)

1760 785 :508

0

[J uU

"'" ~

~ j

:3.0 :3.4

Figure 17. - Effect of altitude pressure on tail-pip~ losses. -1_-

--,..., "-1\ r ..........

~" \ \

\

I I I I

I '.2 90 915

Alt11iado (ttl

-- Sea 1 ... ." ---- 50,,000

(al ~1101' ",-"""tar. tati08 at 00II8'tant Qr,sr-

-- spoed.

5.0 Altitude

(tt) r----Soo.l.oTol. ---- 50,000

r--~ r_

........ ........ ~

1\ ~ \

I

I I I

i

/ / ,.

L ,. I / " V

" /

III 91! 100 105 _1I00I air no., ~ at ",1:04

(bl C __ "1Ol' ",," .. tins 1m...

" V I 1/ ,. V " 1/ /

I V " / ,. /

-,..., ~ r~~~ '1\ / /

", r/~ / ~

I

" V A / " Vi ,. / I

I

I 90 95 1,J)O

(0) 00IIiFN""" .-tor1ot1o. om oporatins l1Doli __ ood.

I V I--r--. /J r-~_ r .... , /' A /

B,. )( \ , ,. / I

I

i I I I

Ll!!

3.( 90 98 100 1~ 90 uX> 105 90 100 10:5

(a) C_"""" .-.-lJrbic. at ~ ..".,..atoc\ opoocI.

~ air nOl" peroeDt ot rated.

(b) ~Na>' O],1On\tins l1DU. (a) OcIIIproumo _tertot1oa om ~t1llS 1JJ>N _~IId. ...

CJ1

1.1

1.0

.9

1.1

1.0

.9

1.1

1.0

AD

A'1 r-a-~ r OA

(a) Iz1i!;I.ne BJ arlal-tloV englno; oorre<rt.ed eJlI!iIle speed, rated.

.r:t. ..... :--- .M

a.

(b) Imgine B, o:dol.-tlDw eng1J)e1 corre<=to<l. ""I!1ne "l'oed, 0.90 of rated.

..n .... n [ oA

IV ~

Fl1Bb.t Mo.eh

- mmber ,

0 0.~7

0 .62 - I:>. :91 OA C

t~ .:.0- ~ L

(ol.) Bnglne F, cantr1:!"usaJ.-tloV eJ1il.ne J corrected omi;1no 1P=ed, rated.

0...

~ !)1T" U ~

rIi

(e) .:!Dg1n<o 1', aontrUupl-flClll' oDj!l.Do; oarreoted ensII>" . .poed, 0.910 of ... toO..

a-L. IA ,.. .r ~

~ ["'[ -c:.

/' ~

o .2 ., .B .8 1.0 0 .2 ., .6 .8 1.0 e R.)'D<IldB _r 1_, ;;;n

(e) IngI.J>o BJ ~-tloV engineJ COZTOCte4 ..,.m. ~, (f) I!I>gIne r, centrl.fu81!l-flQY engineJ .,,,,,:ect.d oq1ne 0.78 of ratoc1. epeeol., O. n of ntoO..

,. 9Zrr:Z

1.1

.9

~~ :>< :<~ 1.1

. ~~ H

1.0

'ilc! ';;~ §

.9 0

1.1

1.Q

.9

!

.M It.. .A I'-'

~ 0 ~ V"

fIl"~

(s.) :Ins!no B, u1al-flow engI:ao, oarl'<Icted ens1no sposd rated "

0 ... I: .err !l. w L

. (b) KDs!no B, s.z:Ial-tlow ona1D>, oarrected ong:I.no opood,

0.90 of rated.

[ - r

-.l~ ~ f~

2326

Flight

f--lI!Ioh

IlIlIIlw'

0 0.S7 f"- a .62

l'J. .91 0

Y ~ ~ L

t:. 0

(d) ~_ r. oan:tr~-!1011 ona1l»; oarreoted e:og.I.no speed, ratod.

r ..A

/ ;>-U <u-

if: (8) Ii:IIS1De ]1', oentr1tupJ.-f'low eng!J1e, "",.....otecl. engine

1l""'11, 0.94 ot raW.

.A

t. r ~ ~

/'

d ~ o .2 ., .6 .S 1.0 0 .2 •• .S .8 1.0

e ~. I\IIIIlber 1l>ieJ:,..,., (0) ~ B, azjal-naw ens1D>; corrected 0IJg:In0I opoed, (t) ~1D> r, oentr1f1lso.l-tlow _!,no; aonoeot&d ens!,no speed,

O.7S ot ratod. 0.71 of ratoll.

J'1sure 21. - 'h-rlation at aotual to pred:Lo"W air flow ratio with h;rnolds muober index.

1.1

1.1

.9 o

•

~

On-...

ro-~ n

( .. ) l!ns1J:l8 B. az1o.l-fi"'" ""Sino; oorrected ong1no o:P""'" , m ted •

r~·u lj>-o... t&-r

'"

(b) bgl". B, u:lal-f'lOII ons_; oorreOted ons- speed, o 90 or rated.

"

'u...... ~ ... ~

.2 .4 .8 .B

J'l.1gIrt -llllIIIbar

0 0.3'1

0 .62 ~ .9J.

.~ n ..J'I J' U ....

(d) Kns- ." oen:trif1l8o.l-floor e~lJ:l8; """",,.ted ~Ino opeed, rated.

! ~ ""

f----

I--

(.) ~ ." ..... tr1ttIgo.l-flOll 8Il!I1J>e; oon-eoted ensuo apoo4, o 94 or mted

I~ ~ ~ '" -'" -L,l

~ 1.0 0 .2 .8 .8 1.0

lle1'lOldII m.I>~r l..tdex , 'II ~

(0) EIlIIIDo B, u1oJ.-fi"", 8!l81no; OOlTtl.ted e~1IIII lIpS""', (r) Ins! ... V, centl'lf1l!!al-fi ... Or1llu",; oorrected ~11111 opeed, 0.78 or mted. 0.71 ot mted.

~ 22. - Tll:ri&t!OIl or aot>.al 1:0 predlote4 t""b_-~let total ~t""" ratio vltb llQnol.4.o Jl1IIIlooO:' 1n4ox •

•

(!D.

.. NACA BM E5lJl5 49

1.3

1.1

.9

1.1

~~~ "' .... 1.5

\.3

1.1

• 9

1.1

1.5

1.3

1.1

.9 0

~~

Fllaht Mach

number . 0 0.37 CI .62 A .91

'\

"'" I\,. ~ N Q

If:.

~ ~ ~ ~

(&) IIlDg1ne B, ax:l.al-rJ.cnr ~1 =ected engil!e speed,· r&ted

Cd) ED&1ne F, centrUui;al.-:C1Dv ~, corrected end.! Ie . .,-peed r&ted.

~ ~

\ \

.

0

~ Po. ~

ro--c t--.. t=¥. d>: ~ I/\. "-.. .....

10 p

erud..Ie ""eeel. 0.90 <4 r&ted. Cel EI:lg1ne F, een~-ficDi" engine, corrected

engine speed, 0.9<10 o~ rated

b'q ~ ~

N N I-

. .6

f"

)

q. \

ra-.~ -A. .~

.s 1 • .0 0 .2 5

Re~ IlIlI:Iber iIldex, ,.jf"

I~ t--

.4.

~ '-£l

.6 .8 1.0

(el :!ngI..ne B, ax:l.al-:rI.ow engine1 corrected (n BngI..rie V, centrUugaJ.-fiair englJ:l.e 1 corrected engine speed, o. 78 o~ rated. engine speed, o.n o.t r&ted..

figure 23. - Variation. o:f &etual to predicted tuel fl.ow ratio with Reynolds lIU!lIber index.

50

2.0

r--

1.6

1.2

.8

2

8

4

° .7

..

01 • NACA RM ESlJl5

Altitw.e (ft)

Pred1cted Fn"1 Sea level ~v ---- 50,OOer

~ Actual F / n

~/ /

Sea-level Fn""", VF Rated speed at 50,000 ft

h '1/

~ /

/

~ V/ ta. t.e<1 spa-ad at

sea level

- V ..,/

. ... -_.

(a) Eosine J.I', oentrifuse.l-f1ow (tl)8ine.

Aclual. F~;""

~~ K Predicted Fn

/. V

.8

V Ra'ted speed at ~O,OOO ft

./ ~sea-lelel F n

~Ra.ted speed L V sea level

.9 l.a._ . Corrected engine speed

Rated BIJgine speed

_,1.1 1.2

(b) Engine B, axial-flow engine.

~ I L

FIgure 24. - Variat10n of _o.orrected net thrust with corrected ensine speed at two ali;1tudeEf. ---,-- :.c.

...

""----"~'

NACA EM ESlJl5 0 __ ._5 •

Eng:!.ne Fllght Mach

number

0 C 0.25 • 0 B .62

~ A .52 G .61.

"q F .67

il 1.0 01/ ----< >--..:::: ~ )..-

.9

..........

~

.8

(a) Ratio of' temperature-limited speed to rated speed.

1..1 -

~ -~ ~ ..............

~ >-- ~ t--... ~

1..0

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

~ .........

~ I

.......... ~ 1"-.. ~~

• 9

.8

(b) Ratio of' mwd.mum thrust to predicted ma:J:!.mr.ml thrust.

}

/ ~ ~~ -~ ~~ I---': .., -fn -

~ 10 000 20000 30 000 , 1 4.0 000 50 000

Altitude, f't

(c.) Ratio ~ actual.-:to predicted specific fuel. eo:OsUmpti·on.

Figure 25. - Variation of actual. to pred:l.cted performance ratios w:l.th altitude f'or several axial.- 8l1d centr1f'uga.1-fiov turbojet engines.

,

51

52 NACA EM E5lJl5

Q) ..-IQ) N..-I N N :3 N

:3 1.1

1.0

.9

1.1

1.0

Fllght Engine Macll number

---0.25 0 C

[J B .62 <> A .52 6. G .61 ~ F .67

-- ~_~~.A)

(a) Ratio 01' maximum thrust with variable-area nozzle to maximum thrust with fixed-area nozzl.e.

...(1)

~> ~

~ I .9

o 10,000 20,000 30,000 _40,000 50,000 Altitude, ft

(b) Ratio o.f specific fuel-c.onsum;ptio~ with variabl.e-area nozzl.e to S'Pecif'ic fu,el. consumption with -fixed-~ea nozzl.e.

Figure 26. - Variation of variabl.e-area nozzle to fixed-area noz­zl.e 'Performance ratios with altitude fqrs~~r~l ~ial- and centrifugal-fl.ow turbojet engines.

•

CO C\J I/') C\J

NACA EM E5lJl5 ----~ ~ a r-I 'H

r-I

t r-I

I ar III O'l

~<d III

1:l8. III O'l 0

tl'al Po~

;'"' a r-I 'H

'"' 1i! od IV .p

fil 1: a t,)

~ III 0

'"' III Po

>: 0 1=1 III .-I 0 .-I

~ III

'"' a III III III

'"' ~ a t,)

.. 120

110

100

90

.

80

70

90

80

p..-: tV

70

60 70

Engine Flight Altitude Mach (ft)

number

0 C 0.25 45,000 0 B • 62 50,000 . ¢ A .52 45,000

~ t:.. G .61 50,000 "I F .67 50,000 /

---r-Tem.perat,ure-lim.1 ted A~ ~-

engine speed I

~ I> I

¢ Ra. ad

I~ spe"';

1<>/ ~ rt

I VI 1/ I

(a) Corrected air f'1.aw.

~led h_ speed

n ~ ~

~ ~ "'" ~ ~ ~ ~ ~ ~ - L.-- .". ""'I

~ ~

80 90 100 110 Correct.ed engine speed, percent of rat.ed speed

(b) Compressor efficiency.

Figure 27. - Comparison of air flow and compressor effioiency characteristics for several engines.

•

~ I"

laO

54

.p m

~ rei II) Q)

~ .p

~ rei II) .p

j f .p

a~ rei gj

~ ~ .,... rei .p II) 8

r-I

~ i ~ II) I

~S C,)

AM&!a NACA EM E5lJ1.5

I

~

Eogine Flight Mach

number 0 C 0.25 Cl B- .62

1.8 <>- A .52 .t.. G .61 ~ F .67

1.6 ) ~

~~

1.4 # V V

/

/ y V ./ ~

1.2 / C0 ~ ~ ~ ~ V

1.0 ~ ~

(a) Ratio of corrected predicted ma XiDDl]!1 thrust to sea-level

1.2

1.0

.8 o

rated thrust. --

....(~ ~

.... ~-=:::; ..... ~

II -- - - -

10,000 20,000- 30,000 Altitu:le, ft

(b) Batio of' predicted specific fuel Co,p,Bumption to sea-level specific-fuel consumption.

Figure 28. - Variation of predictedto sea_-level pe:rformance ratios with al tltude f'or seVeral ax1al- ani lyerttrifugal"':fl,oW- turbojet - engInes.

•

rn -- N m

, ..

NACA BM E5lJ15

1.8

1.0

1.2

1.0

-

Engine Flight Mach

Dlllllber

0 C 0.25 C B .62 <> A .52 h. G .61 ""I F .67

~ I' --.... t'" ............ i""" / 'r---~

/ L

17

~ r-I)

/ V ~ V V -~> ....;

/ y

~ ~

~ lP ~

(a) Ratio of' corrected maximum. thrust i;o sea-level rated thrust.

..-f

~

/~

!r--'" V V .-l. 1;

~ r-- ~ V r-~

~ r .8

° 10,000 20,000 30,000 40,000 50,000 Al.titude, f't

(b) Ratio of' actual epecif'ic fuel consumption to sea-level spec1f'ic fuel consumption.

Figure 29. - Variation of' actual to sea-level performance ratios with altitude for several a:x:lal.- and centrifugal-flow turbojet engines.

55

56

Engine Flight Altitude Maoh (ft) . ,

·mnnber

0 C 0.25 45,000 ~ C B .62 50,000 <> A .52 45;000

~ V A G .61 ~O;OOO ~ F .61 50,000

·llP V I>

~ ~ .d ~

~ V""" C1

.... -(a) Engine pressure ratio.

~

~ Vp

l0 ~ ,/

V ~ ~

~ V 0'" « LV'

80 90 . 100 _ Corrected engine speed, peroent of rated speed

(b) Engine temperature ratio •.

NACA EM E5lJl5

~ 120

Figure :30. - Variation of e:oe:ine ·:pl"e,ssUl'E! ratio_ !Wi.temp~z:~tUFe.r8.tio with corrected engine S1)eed f'or BErVera1 arial- arid centr1f'uga,l-:f1ow· engines. . -. .

NACA-Lang!ey - 12-:11-&1 - 471

. ·t

........ --I

--.. ---.--.--------~~

, 1111~l~I;llliilrnjl~~l~l]lllllll~ 3 117601435 1630

1

I

I

I