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RESEARCH MEMORANDUM

ALTITUDE -CHAMBER PERFORMANCE OF BRITISH

ROLLS-ROYCE NENE II ENGINE

III - 18.00-INCH-DIAMETER JET NOZZLE By Ralph E. Grey, Virginia L. Brightwell, aad Zelrnar Barson

Lewis Flight Propulsion Laboratory Cleveland, Ohio

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

https://ntrs.nasa.gov/search.jsp?R=19930086320 2020-06-17T12:41:58+00:00Z

. HATIOIKL AsrvIsoRY C O W C T E E : E'OR AwOmAzrrICS

A L T I T U D E - C a PERFORMANCE OF BRITISH

ROILS-ROYCE NEloE IS ENGINE

An altitude-chamber investigation was conducted a t the WCA ~ e w i s hboratory to determine the a l t i tude performance character- i s t f c s of the Britfsh Rolls-Rope Nene 11 turkojet =#.ne eth an 18.00-inch-dLameter j e t nozzle. Results m e presanted far sFmu- lated &l.tftudes~frcan sea level to 65,OOO f e e t and for ram pressure ra t ios froml.10 t o 3.50 (corresponding t o flight Mach numbers from 0.37 to 1.47, a s s a g 100-percent ram pressure recavery).

Typical performance-data plots are pre.sented to show graph- ically the effects of a l t i tude and of fltght ram pressure ratio.. Conventional correction methods were applied to the da.b to deter- mine t h e possibil i ty of generalizing each perfollpaslce pazameter t o a single curve. A complete tabulation of corrected and uncorrected enghe-perforrtlance p a a m s t e r a is gresented. A compariscm of engFne p e r f o m c e with the 18.75-, 18.41-, and 18.00-Fnch-diameter j e t nozzles and without a j e t nozzle is made t o show the effect of changes i n nozzle s ize under simulated-flight cmditions.

The Investigation showed that engine perfommce obtalned at anY one altitude could not be used t o predict performace a t other al t i tudes above 30,000 feet with the 18.00-inch-diameter j e t nozzle. For varying ram pressure ratios at a given a l t i tude , engine per- formance can be predicted *am data representing other ram pressme ratios only when c r i t i c a l flaw exis ts in the jet nozzle,

A cornpaison of the .engine performance with t h e th ree jet- nozzle sizes and without a jet nozzle a t an a l t i tude of 30,000 f e e t and a ram pressure ratio of 1.70 indicated that the 18.00-fnch- diameter J e t nozzle gave the lowest values of net-thrust specific fue l consumption a t practically all engine speeds. A t lower ram pressure ratios, the 18.4l-inch-diameter jet nozzle gave lower

2 NACA RM ESOA31

values of ne t - tMst spec i f ic fuel consumption at high engine speeds. J e t thrust, net thruet, fuel cansumption, and t a i l - p i p e indicated gas temperature generally increased with use of the smsller nozzles.

The al t i tude performance investigation of a British Rolls- Royce nene II 'engine vas conducted in an altitude chamber at the NclcA Lewis laboratory durlng 1948. Three dffferent jet-nozzle dimetere were used in the investigation of this engine to deter- mine the effect of nozzle size cm enghe performance. A " b e d amount of d a t a was obtained without a j e t nozzle attached to engine tail pipe.

The principal objectives of the Investigation wwe t o deter- mine ' the altitude performance with a 18.00-inch-diameter jet nozzle and t o determine t he range of simulated-flight conditions over which the performance parameters mfght be generalized t o a single curve. The effect of change in jet-nozzle size an al t i tude performance was of interest, particularly with reference'to enghe specific fuel consumption, because a j e t nozzle amaller than stand- ard can be used at cruise cmditions without exceeding allowable temperatures. A emaller jet nozzle should. give higher thrust and. possibly lower epeclfic fuel canstmrp.t;ian over nearly the entire range of angine speed.

The effects of a l t i tude and f l i gh t speed on tlie over-all engine performence using t h e stauda,rd 18.75- and a n 18.41-inch- diameter j e t nozzle are presented in references 1 and 2, respec- tively. The mer& engine performance wing an 18.00-inch- diameter Jet nozzle i s presented herein. Results are presented for slmulated-flight conditions varying in altitude from sea level t o 65,000 feet and III ram pressure ratio frai L l 0 - b 3.50. These ram pressure ra t ios carreapand t o f l i g h t Mch number8 from 0.37 to 1.47, a s s e e ; 100-percent ram pressure recovery. The conven- tional method of reducing data t o am-level conditions (reference 3) was used to determfne whether perfornmce could be generalized; that le, whether data obtained at ORB altitude and ram pressure r a t io can be used t o predict performme at other c~nditi~ne of a l t i tude and ram pressure ratio. A comparison of' performance w i t h three jet-nozzle sizes and without a jet nozzle (open t a i l p i p e ) i e presented a s an indication of engine performance with a variable- area j e t nozzle; generalizatian of performmce with varying jet- nozzle area, however, i s not included..

. NACA RM E5OA31 3

A cutaway v i e w of the British Rolls-Royce m a n e 11 power plant, which is a through-flow tmbojet angine having nine combustinn chambers, is darn in figure 1. The engine incorporates a single- stage double-entry centrifugd. compressor (tip diameter, 28.80 in. ) driven by a single-stage reaction turbine (tip diameter, 24.53 in.). The turbine-nozzle area is 126 square inches and the standard jet- nozzle area is 276 muare inches. The dry engine weight is approxi- mately 1720 pounds (starting panel and generator included) ; the rceximum diasleter (cold) is 49.50 inches, giving an effective frontal area of 13.36 square feet . The ses-level e@ne perforoBnce with the standard 18.75-inch-diamster jet nozzle (reference 4), based on Rolls-Royce static test-bed data, is:

J e t thmrst Specific fuel mine speed 1 consmptian (rPm)

(lb/(hr) (lb -st) 1 Take -off rnlitary

1.04 12,250 .5000

”” 2,600 120 Idle 1.02 U,5# 4000 Max. cruise 1.04 12,250 5000

From these values it can be seen that t h e rated military thrust per unit weight of engine is 2.91 pounds thrust per pound weight , and the rated military thrust per unit of frontal a m a is 374 pounds thrust per squaze foot. The mxinunn allowable tail- cane &as temgerature is 1365O F with the standard 18.75-inch- diameter jet nozzle.

A sea-level acceptance run of the engine with the standard 18.75-inch-diameter jet nozzle, with m i n i m u m research instru- mentation installed, showed a thrust of 5110 pounds and a spe- cific f u e l consumption of 1.01 powda per hour per pound of thrust at an ewine speed of 12,261 rpm.

Altitude Test Chamber

The engine was installed in an altitude test chmiber 10 feet in diameter and 60 feet loa@; (schemticdJy shown in fig. 2). The inlet section of the chamber (surrounding the engine) was separated f’rom the exhaust section by a steel bulkhead; the engine tail pipe passed through the buU&ead by means of a law-friction seal. The

4 NACA RM E5aA31

seal was composed of three floating asbestoa-board rings so mounted on the tail pipe as t o al low thermal expansion in b o a r ad ia l and axial directiana, as well as a reasonable amount of lateral movement to prevent bfnding.

3hglne thrust U&B mea~ured by a balanced-pressure-diaphragm- ty-pe thrwt indicatar outside the t e s t chamber, cc-rmnected by a linkage to t h e frame on w h i c h tf ie engine was mounted in the chamber.

An A.S.M.E. t n e f la t -plate or i f ice mounted in a straight ruzl of 42-inch-diameter pipe a t the apprpach to the t e s t chamber was provided f o r measuring engine air cmsmptim. Because of the large variatim in atmospheric conditions investigated, considerable difficulty'was encountered with condensation i n the orif ice differential-pressure lines despite repeated attempte t o remedy this si tuation. Engine air commption 'bl&~ therefore calculated from pressure andl temperature rueaeurements in t h e tail pipe, as described in the appendix.

Ram-air pressure was controlled by a main, electrically operated butterfly valve in the 42-Fnch air-supply line, bypassed by a 12-inch, pneumatically operated V-port valve. A i r WEB supplied by ei ther a combustion-air (moist, room temperature) syetem ar a refrigerated-air (dry, cool&) system at temperatures near those desired* Mnal cantzol of air temperature was accom- plished by a set of electric heaters in the bypass line immediately preceding the entrance to the test chmiber. Tbe air entered the test chamber, passed through a set of strraightening vanes, and then entered the engine cowl. The purpose of the cowl was to prevent direct circulation of heated air f r o r n t he region of the tail pipe & combustion chambers i n t o the aft inlet of the c~mpessor . The air so heated was therefore mixed with the cooler air supply before enterkg the compreasor.

The exhaust jet was discharged Into a diffusing elbow mmtsd. i n the exhaust section of t h e chamber. This elbow ducted the gpsee into a dry-type pr- cooler. Control of the exhaust preesure was obtained by a main, electricaUy operated butterfly valve, bypassed by 8 20-inch, pneumatically operated butterfly valve. The gases then passed through a dry-type seconbzy cooler and thence into the system exhausters.

M tn -4 C U

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NACA RM E5QA31 5

Instrumentation

Compressor-inlet temperature and total pressure were m e a s u r e d by eight probes, each comprising an irm-ccmstantan thermocouple aSa a total-pressure tube. Four probe8 weze equally spaced azound the periphery of the front campressor-inlet s c r e m and four around the back screen (station 2, fig. 3). Cmtzol of r a m pressure and temperature was based on t h e averaged readings of t h e eight probes. Campressor-discharge pressures were measured at the exit of compressor-discharge e l b m 1, 4, m d 7 by 8even total-pressure tubes in each elbow.

Engine tail-pipe temperatures at station 6 were measured by 25 c h r o m s l ~ ~ ~ , stagnation-type thermocouples lo-eea in an instrument ring. The 5astrment ring also inclrded 24 total- pressure probes, 14 static-pressure probes, and 4 wall static- pressure taps. This instrumentation was located. approxjmately 18 inches downstream of t h e tail cone. In addition, t h e four standard nene e n g i ~ e tail-cone thermocouples supplied by Rolls- Royce Ltd. were mounted in the tail cane and were used for engine- control purposes. Pressure a& temperature instztnnmtaticdl was also located at other statim13 throughout the enghe; measurements from +&is hstrmntatim are not reported.

AU. peasures, including the thrust-indicator-diaphraepl pres- sure , were instantaneously recorded by photographing the nanometer panelo Temperatures were recorded by two self-balancing, scanning potentiometers, which required about 3 minutes to record all en- temperatures.

-gin8 speed was measured by an impulse counter, which operafed on the frequency of a t;hree-phase generator mounted 0 ~ 1 the accessory case of the engFne. Actions of the counter and a timm were synchronized.

Fuel consumption WBS measured by a calibrated variable-area- orifice flow meter, &ich alzowed full-scale read- for various ranges of fuel flow by changing the orif ice flow area.

With the exception of air condtrmption, performance data were generally reproducible within 2 percent. Air-conslrmption d a t a scattered appreciably at hfgh engine speeds-and were, in general, reproducible only to within 5 percent with a few points showing even greater scatter.

6

Procedure

mACA RM E50A31

Performance characteristics of t h e engine were determined over a range of engine speeds at sirm;rlated altitudes from sea level to 65,000 feet and. ram pressure ratios from 1.10 to 3.50. Inlet-air temperatures wre, in general, held to within 3' F of NACA standard d u e s corresponding to the einmhted-altitude and ram-pressure- ratio conditions. Comgressor-inlet total pressures were held at values correeponding to the simlated-flight conditions, assumfng 100-percent ram pressure recovery.

A surmrrary of yerformance and operatioaal data obtained at 8i?IlUkted-dtitUde conditions is preeented in table 1. Altitude data corrected for mdJ. variaticws in conrpressor-inlet pressure and temperature settings and for variation@ in exhaust-pressure settings a r e s w i z e d in table II. Table IS also ilncludes the d a t a corrected to conditians of mACA standard sea-level static pressure and temperature at the ocmrpreseor W e t .

Simulated. -Flight Performance

Effect of altitude. - mica1 performme data. from table 11, obtained at a ram pressure r a t i o of 1.30 and aimulated altitudes *om sea level to 60,000 feet, a r e presented to show the effect of altitude on jet thrust, net thrust, air consumption (cooling air exoluded), fuel cansumption, net-thrust specific fuel consumption, and tail-pipe indicated gas temperature (fige. 4 to 9, respectively). The trends shown &re similar to those discussed in reference 1; that is, Jet thrust, net thrust (except at low engine speeds), air con- sumption, and f u e l canerumption rapidly decrease with an increase i n altitude and net-thrust apecific fuel consungtion generally decreases up to an altitude of approximately 30,000 feet, above which this trend revereee'to give higher specific fuel consumptian at higher altitudes. Although the data plotted for an altitude of 60,000 feet axe too scattered to indicate this reversal conclusively (fig. 8), plots (not included herein) of other data from table I1 make the reversal in trend evident. This reversal, discussed in reference 1, is a result of decreasing inlet-air temperature, which increases the conpressor-tip Mach number, thus producing an increase in the com- pressor preseure ratio and cycle efficiency. The reversal therefore apparently .takes place at the tropopause (35,332 ft based on NACA standard atamsphere).

.

NACA RM E50A31 7

The specific-fuel-consmrgtion curves are ccmputed from values obtained f r o m the faired fuel-consutqtion and net-thrust curves; any discrepancies that occur between the fuel-cansumptian and net-thrust data and the faired curves are carried over to the specific-fuel- consmptiun curves. The actual data points therefore in m a y cases do not fall on the computed c w e .

At engine speeds below 10,000 rpm, tail-pipe indicated gas temperature (fig. 9) decrea8ed as altitude was increased to the tropopause and then remained constant with further increase in altitude. A t engine speeds above 10,000 rpm, this trend was reversed. This reversal in trend takes place at engine speeds lower than 10,000 rpm for ram pressure ratios greater than the sample 1.30 data and at higher engine speeds for lower ram pres- sure ratios.

’ Effect of ram pressure ratio. - Perf oraEtIlce data obtained at R simulated altitude of 30,000 feet and at r a m pressure ratios from 1.10 t o 3-00 are presented- to show .the effect of ram presswe ratio an jet thrust, net Ghrust, air cansumption, fuel cmstrmption, net- thrust specific f u e l consumptian, and t a i l -p ipe indicated gas temper- ature (figs. 10 to E, respective*).

The lncrease in air density at the engine inlet that acccapn3es an increaee in ra~n pessure mt€o generally increases j e t thrust, air cmsum-ption, and fuel consmption throqjktout the range of engine speeds FnvestZgated- Net thrust increases with increasing r a m pressure ratio at high engFne speeds but decreases with increasing ram pressure ratio at low engine speeds. For the sample data shown (altitude of 3O,,ooO ft), the reversal in trend occurs at a p p r o ~ t e l y 1 0 , o O O rpm. Net-thrust specific fuel. consumption increases with increasing ram pressure ratio. The tail-pipe indicated gas temperature in general decreases slightly with increasing ram pressure ratio. T M e decrease is mall and somewwt incornistent and could be interpreted &a data scatter at the higher engine speeds. As would be expected, an appreciable decrease in temperature occura at the lower englne speede where there is a tendenog for the e w h e to w3MJIlil.l.

These trends with varying ram pressure r a t i o &re simtlar to those discussed in greater d e t a i l In reference 1.

Generalized Performance

PerfoMnance d a t a * representing engine operation at altitudes fram sea level to 65,000 feet and at ram pressure ratios f’rm

8 NACA €&I E50A31

1.10 to 3.50 were reduced in the conventional manner (reference 3) to NACA etandard sea-level conditions. The development of this method of generalizing data involves the concept of flow similar- ity and the application of dimensional analysis to the performance of turbojet engines. In this development, the efficiencies of engine components are considered to be unaffected by changes in flight oanditions.at a given corrected engine speed.

Effect of altitude. - Ty-pical corrected engine performance d s t a (table 11) obtained at a ram pressure ratio of 1.30 and sim- ulated altitudes from sea level to 60,000 feet are compared to show the effect of altitude on the corrected values of Jet thrust, net thrust, air consumption, fuel consumption, net-thrust specific fuel consumption, and tail-pipe indicated gas temperature (figs. 16 to 21, respectively).

The corrected values of Jet thrust and net thrust (figs. 16 and 17) generalize for all altitudes up to 30,000 feet. At higher alt i- tudes, the tbruet decreases with increase in altitude. This decrease in thrust.with-altitude is less than that shown in references 1 and 2 because with the 18.00-inch-diameter jet nozzle the engine operates at higher pressure and temperature levels. The decrease in air den- sity with increase in altitude therefore has less effect. Also, because an appreciable matter exists in the-available data for high altftudes, no consistent trend in air consumption with increase in altitude is indicated (fig. 18). Corrected fuol coneumption (fig. 19) increases slightly with increase in altitude at high values of corrected speed. Plots of other data f rom table I1 show that at low engine speeds, fuel consumption .hcrea.see rapidly wfth increase in altitude, as is also sham in references 1 and 2. The corrected net-thrust apecific fuel consumption curves (fig. 20) generalize up to an altitude of 20,000 feet. Above 20,000 feet, the corrected specific fuel consumption increases with increase.in altitude. The 60,000-foot data points do not fall on the Codputed curve, as explained in the discussion of figure 8 . The corrected tail-pipe indicated gas temperature (fig. 21) generalizes f o r a l l altitudes investigated .

Effect of ram pressure ratio. - The conventional method of generalfzing data was specifically developed to &just for changes in the pressure and the temperature of the atmosphere In which the engine ia submerged. .A variation in ram pressure ratio (flight speed) changes the performance characterigtice because i t - . h a e the effect of changing the campressiqn ratio of the engine. In general, the increase in operating pyrssure-that accompanies increase in ram pressure ratio raises the total expansion pressure ratio of the enghe (frog turbine inlet-to jet-nozgle throat) until critical flow

NACA RM E50A31 9

is established in the jet nozzle. After critical flow is estab- lished, the expansion pressure ratio of the engine remains constant with further increase in ram pressure ratio. The engine is then effectively submerged in an atmosphere having.a static pressure equal to the pressure existing in the jet-nozzle throat and is operating at a constant effective ram pressure ratio. The effective ram pressure ratio is then equal t o the ratio of the compressor-inlet total pressure to the Jet-nozzle-throat static pressure. With criti- cal flow in the jet nozzle, generalization of flow characteristics within the engine should be possible within the limitations discussed in connection with altitude effects.

Typical performance data obtained at a simulated altitude of 30,000 feet and ram pressure ratios from 1.10 t o 3.00 are compared to show the effect of ram pressure ratio on the corrected values of jet thrust, jet-thrust pazmeter 'J + '$7, net twst, air can- sumption, fuel consumption, net-thrust specific fuel consumption, 8n.d tail-pipe indicated gas temperature (figs. 22 to 27, respectively).

Corrected jet thruet (fig. 22(a)) does not generalize but cor- rected jet-thrust parameter (fig. 22(b)), for which the development is given in reference 1, generalizes for all conditions for which the jet nozzle is choked. The corrected net thrust of figure 23 appears to generalize for ram pressure ratios greater than 1.30 at the higher speeds, but data for ram pressure ratios less than 1.30 do not generalize. Inasmuch as net thrust is a f'unction of jet t h r u s t and air consumption and jet thrust did not generalize, there is no reason t o expect net thrust to generalize. At higher flight speeds (ram preesure ratios), however, the momeRtum of the lncaming afr is greater for a given mss flow; this larger quantity, when subtracted f'rm the higher jet-thrust values of figure 22 (a) , causes the corrected net thrust to generalize for ram pressure ratios above 1.30. Corrected air consumption (fig. 24) apparently generalizes at a l l ram pressure ratios. Corrected fuel consumgtion generalizes at the high engine speeds when critical flow exists in the jet nozzle (fig. 25). At lower engine speeds the fuel consumption decreases with increase in ram pressure ratio. Net-thrust specific fuel c m - sumptian (fig. 26) shows reasanable generalization for ram pressure ratios of 1.30 and above. Data for a ram pressure ratio of 1.10 show slightly l m r vslues of specific fuel consumption. The tail- pipe indicated gas temperature (fig. 27) also generalizes to a sin- gle curve for engine speeds at which critical flow existed in the jet nozzle. At lower engine speedls the corrected tafl-pipe indi- cated gas temperature decreases with increase in ram pressure ratio.

6

10 NACA €44 E50A31

Effect of Jet-Nozzle Area on Performance

The perf'omnance of the engine using an 18.00-inch-diameter jet nozzle is cozpared at an altitude of 30,000 feet and a ram pressure ratio of 1.70 with performance using: (a) the standard 18.75-inch- diameter jet nozzle (reference 1); (b) asl 16.41-inch-diameter Jet nozzle (reference 2); and (c) engine tail pipe without a jet nozzle. The tail-pipe diameter is 22 inches; therefore, the data without a Jet nozzle will be referred to as the "22-inch nozzle data." These results are presented In figures 28 to 38. The changes in perform- m c e cawed by changeB in jet-nozzle area follow the expected trends diacussed in reference 2. Tail-pipe total pressure increases with decrease in nozzle size (fig. 28). Compressor pressure ratio at a given engine speed remains nearly constant fw t-he-range of air flaw encountered in this investigation (fig. 29) except at the lower engine speeds. Becatme tuibine-inlet pressure remains nearly con- stant and tail-pipe total pressure increases with,decreaeing jet- nozzle area, total-pressure ratio across the turbke decreases with a decrease in jet-nozzle ares (fig. 30). In order to maintain the required cmpressor work per pound of air, which is independent of nozzle size (fig. 31) and represents nearly the entire turbine power output, it is necessary to operate the turbine at a higher temper- ature level as the Jet area is decreased, which results in an increase in both turbine-inlet total temperature (fig. 32) and tail- pipe temperature (fig. 33). Except at the high engine epeeds, the increases in turbke-inlet total temperature and in tail-pipe indi- cated gas temperature a m nearly equal.

For critical flaw in the turbine nozzles, a i r flow is essen- tislly proportional to the turbine-inlet pres~we, which is nearly canstant, and invereely proportioagbl to the square root of turbine- inlet temperature, which increases with a decrease in nozzle size; therefore, air consumption decreases with decreasing jet-nozzle area (fig. 34). Because the air-consumption data included herein, as well as those of references 1 and 2, were not efliciently con- sistent to indfcate trends of small magnitude, the curves of ffg- ure 34 were obtained from a sFngle faired curve for each nozzle size, using corrected data for altitudes up to 30,000 feet at a ram pressure ratio of 1.70. Air consumption generalizes with alti- tude up to 30,000 feet, as is sham in figure 18; it i a therefore possible to-invert the correction factors and to apply them to the corrected parameters to obtain a smooth curve of proper magnitude through the a.ctual altttude data points. The air-consumptim values of figure 29 were obtained from these same fatred curves. The data pohts of figure 34 are actual altitude data, and when plotted alone, they do not indicate the trend too clearly. The expected increase in f u e l consumption accmpaniee a decrease in

P

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w

EIACA RM E50A31 11

nozzle size (fig. 35). Jet thrust increases with a decrease in nozzle size (fig. 36) because of the increase Fn tail-pipe total pressure. The trend followed by net thrust (fig. 37) is similar to that for jet thrust. At a ram pressure ratio of 1.70, net- thrust specific fuel consumption (fig. 38) decreases with decrease in nozzle area at most engine speeds; at high engine speeds, the three smaller nozzle sizes give similar values of net-thrust spe- cific fuel consumption, whereas the 22-inch nozzle gives a much higher value. At lower ram pressure ratios (flight speeds), how- ever, the 18.41-inch-diameter jet nozzle gives the lowest value of net-thrust specffic fuel consumption omr a larger portfon of the high-engine-speed rage.

The follouing results were obtctked from an altitude-chamber investigation of the performance of a Britfsh Rolls-Royce Nene I1 turbojet engine using an 18.00-inch-diameter Jet nozzle:

1. ‘Engine-performance pasameters, except for air consumption and tail-pipe indicated gas temperature, could not be predicted f o r alti- tudes above 30,000 feet from data obtained at one particular altitude.

2. Performance data at any ram pressure ratio for which critical flow existed in the jet nozzle could be used to predict performance at any other ran pressure ratio in the critical flar range within the limits of this investigation.

3. The 18.00-inch-diameter jet nozzle indicated 8 lower value of net-thrust specffic fuel urnsumptian over substantially the entire range of e@ne speed investigated than either an 18.41- or the standard 18.75-inch-diameter jet nozzle at an altitude of 30,000 feet and a ram pressure ratio of 1.70. At a r a m pressure ratio of 1.30, however, the 18.41-inch-diameter jet nozzle gave the lowest values at high engine speeds. The engine operating without a jet nozzle attached to the tail pfpe gave a much higher value of net-thrust specific fuel consumption. Jet thrust, fuel consumption, and tail- pipe indicated gas temperature a l l increased when Bnaaller Jet noz- zles were used, whereas air consumption showed a slight debrease.

Lewis Flight Propulsion Laboratory,

Clevelmd, Ohio. National Advisory Committee for Aeronautics,

12

m m I x - CLILCTIIATIONS Symbols

HACA RM E5aA3l

A

D

F

g

H

J

I(:

M

R

P

P

R

T

t

V

wa

Wf

wg

Y

area, ag f t diameter, f t

thrust, lb

acceleratim due t o gmvity, 32.2 ft /sec2

en-py, Btu/lb

mechanical equivalent of heat, 778 f t - lb/Btu

thrust c"t

Wch nmiber

=*e speed, rpm

absolute t o t a l pressure, lb/sq f t

absolute s t a t i c preseure, lb/sq f t

ss c a s t s s t , 53.3 ft-lb/(lb)(oP)

total temgerature, 91 s t a t i c tmperature, 91 velocity, ft/sec

air consumption, lb/seo

fuel consumption, lb/&

g3s floW, lb/8eC

r a t i o of specific heats

NACA RM E5QA31 13

8 ratio of compressor-inlet absolute to ta l pressure to absolute static pressure of NACA standasd atmosphsre at sea level

8 ratio of campressor-inlet absolute total temperature to absolute static temperature of EIACA standasd atmosphere at sea level

Subscripts : *.

c conrgressor

d thrust-measurfng diaphragm '

n net

p airplane

s seal

0 free stream

2 compressor inlet

3 ccmrpressor discharge

4 turbine W e t (canbustion-chamber discharge)

5 tail cone (turbtne discharge)

14 WlCA RM E5oA31

Methods of Calculation

Thrust. - Thrust was oalculated by adding to the indicated t h r u e t i n e d from t h e altitude-chamber ' thrust indicator) a correction factor accounting for the pressure differential across the tail-pipe seal. The re lat ian used was

FJ =I Fi + A,& - Po) where

and the seal area

A i r omsumption. - Engine air consumption wa5 calculated from meamrements of temperature and t o t a l and s t a t i c pressure in the tail pipe. Total-pressure profiles across the tail pipe were'plotted fo r each data point; the profiles were then read st eight points, so selected as t o divide the tail-pipe area into four equal cmcen- t r ic , annular areas The fo l lming formula was then applied t o each of the four areas z

where

A 1/4 x tail-pipe area (cold) . . . .. ." ..

AH enthalpy difference between total- and static-pressure condi- t ime , determined from reference 5

The s t a t i c temperatwe in the formula. waa calculated from the indicated temperature by

NclCa EM E5OA31 15

where the temperature ra t io was determFned from %he tai l-pipe t o t a l t o s t a t i c pressure ratio by means of reference 5. The factor 0.8 is the selected average value of thermocouple recovery factor based on instrument calibrations.

Engine air cansumption uas then determbed from the follaring re lat ion by adding the gas flows through the four annular areas and subtracting the fuel flow:

W f w* = wg - - 3600 Simulated flight speed. - The simulated flight s p e d at which

the engine was operated was determined from

where y was assumed t o be 1.40.

Net thrus t . - Net thrust was calculated frm jet thrust by subtracting the momentum of the free-stream air approaching the engpne inlet , according t o the relation

where V is simulated flight speed. p

Fli'&k Mach number. - Flight Mach number wa8 calculated from the compressor-inlet total pressure, assuming 100-percent ram pressure recovery

where 7 was assumed t o be 1.40.

16 NACA Izpi E5OA31

1. B&rson, Zelmas, and Wilsted, E. D.: Altitude-Chamber Performance of British Rolle-Royce Nene I1 Engine. I - Standard 18.75-Inch-Diameter Jet Nozzle. RACA EM E9123, 1949.

2. Armetrong, J. C., Wilsted, E. D., ardl Vincent, IC. R.: Altitude- Chsmber 'Performance of British Rolls-Rope Nene I1 Engine. II - 18.41-hch-Diameter Jet Nozzle. RACA IM E9I.27, 1949.

3. B o l l a y , William: Performance Variation of Gas Turbines and Jet Propulsion Units with Atmospheric Inlet Conditiana. Pqwer Plant Memo. No. 1, Bur. Aero., Navy Dept., March 3, 1943.

4. Anon.1 Rolls-Royce Nene II Performance and Installation Data. TSD 57, Dec. 1947.

5. Amorosi, A. : Gas Turbine Gas Chmts. Res. Memo No. 6-44 (Navships 250-330-6), Bur. Ships, N a v y Dept . , Dec. 1944.

NACA RM E50A31

i

18 NACA RM E50A31

-

j - - 1

3 4 8

6 7 8 9

10 11 la

14

l6 16

18 17

20 19

21 22 23 24 86

26 27 28 eo 50 31

55

34

38 35

37

la

sa

a8

39 40 41 42 43 44

46 48

48 47

48

W 61 ca 63 M 66 Ea 87 68 -

-

ma

i

g< 9

+I

& - X)-ine 62.87 64.M

do.W

70.37 6 7 . S

57. sa 71.28 79.47 8S.W

6l.W 6e.m 74 .w 66.03

78.87

26.03 55.78

(8.00 44.41 60.77 50.98

81.14 e.4.30 80.80

77.46 74.34

24.58 28.6.2 29.11 51.14

27.6s SLL8 34.m 56-87

54.29

36.57 31.27

SB. 78 37.49 24 .M 30.44

42.14 U .40 44.14

(2.04 44.34 46.61 46.07 48.88

m.64

37. io 42.67 41.28

46.28 48.1.E

30.- 46.68 61.30 64.52

1 ?

I I

I

E I

I I I I I

:

3

3

a 3

S

a a 3

a a S a a a 3 a a -

- a0.a 28.84

29.81

89 .as 90.84

W.6S W.M

EO.65 20.36 20.46 02.59 80.66

w.7c w.Bc 15.7c 15.76 15.66 w.7c w.79 w.74

w.04 W.69

15.54 8.66 8.81 8.76 8.68

8.76 8.76 8.65 8.66

8.M 8.01 8.81 8.74 8.78

8.83 8.00 8.81 8.88 9.05 8.88

8.74 8.W

8 -74 8.74 8.74

8.69 8.48 8.64

8.78 8.78

8.68 8.64 8.79 8.99

a8.a

2o.w

- 38-13 58-76

38.91 38.91 38.74

s4.10 s4.01 s4 .09 S I . 14

36.01 36.14 S5.w S5.04 a4.88

17.47 17.58

17.61 17.42

17.45 17.46

30.04 SO.04 29.w 2e.DB 08.04

9.66 8.99 9.86 9.80

u . 2 0 l l .38 11.74 ll.68

w.4a 15. 4T m.4a w.40 w .e7 ls.05 l5.14 15-46 16.lS 36.22 15. 18

w.04 16.14 L6.14 16-24 18.14

17.71 17.85 1 7 . a 17.76 17.61

20.37 20.U

20.44 a0.49

- 1.29 1.50

1.30 1.30 1.m 1.86

1.88 1-68

1.68

1.70 1.78 1.71 1.70 1.70

Le8 1.28

1.26 1.27

1.28 1.27

0 . l a 2.19 8.19 8 . U e.81

1.l2 1.13 1.10 1.14

1.29 1.30 1.56 1.58

1.52 1.68 1.52 1.63 I .81

1.m 1.W 1.75 1.70 1.68 1.71

1.70 1.75

1.74 1 . 7 3

2 . 0 4 2.06 2 . 0 4 2 . 0 8 2 . 0 0

2.34 Z.S6 2 . 3 5 2.27

1 . m

-

- l4 14

la la w 14

14 14

l4

eo W eo eo 22

W

l.E lE

w ls 15 I3 14

21 21 21 41

20 20 80 W

ls ld 16

15 ld

aa e2 w 82 22 82

m

ea m

20 m

eo 20 20

l5 ls l8 16 I3

la l5 la 16

105

Bo

184

m

ffl

149

1W

146

56

15a

148

14 4

145 1.461 4 w 1.570 6 l 6 2.128 700 2.837 875 2.960

485 1.489 600 2.m

1m S.M9 8w 3.m

la75 4.3w wB0 4 . u 4 440 1.758 800 2.678 a w s.m

l a 5 4.001 L160 3.088

NACA RM EWA3l 19

i E.lW 16 2.7- 16 5.164 14 5.679 16 4 . W 16 - 145 m

1m

146

66

68

146

67

m

161

147

169,

40

168

18S

ea

- e4.m e4.w

e4.02 84.18

m.97

as.02 26.01

26.84

7.50

7.16

7.86

9.m 9.68

9 .57 9.47

e5.94

7.20

7.90

u.m

t1.m

11.20 11.m

L1.14 11.-

L9.86 LQ.16 L9.17

6.91

5.88 6.85

7.56 7.58 7.58 '1.56

11.94 11 .99 18.05 u.9s P.76 8.88 e . n 5.79 s.7a 5.62

4.91

7.6s 7.- 7.45

S.48

8.86

4.05

- 8.79

8.74 6.74

8.6s 8.79

8.89 8.88 8.10 9.04

6.56 6.46

6.m 6.46 6-46

6.45 8-80 6.66 6.55

5.w 6.60 6.66 5.60 6.60 6.70

6.89 5.54 6.79

5.26 5.50 5.S

5.40 5.46 5.40 5"

5.79 S.69 3.44 4.m

8.M 1.96 1.-

8-05 8.05 8.00

1.96

2.04 2.14

8.80

1.66

1.56

1.86 -

1.660 19 1.668 19 s.m 19 4.348 34

8.457 90

5.979 80 4.454 PO 4.605 80

s.881 eo 49 40.000 68 40.000 70 40,030 78 40.000 71 40.000

8.4W 16 5 . m m 3.887 Po 4 . 4 ~ Po

1.780 es 1.97s p6 5.126 95 S.8W 16 4.190 16 4.- 16

1.700 lE 8.661 18 s.u7 la

5.161 SO s.888 eo 4.sm eo

P TI 40,ooo 79 40.000

81 40,000 eo 40.000 88 40.000 78 40,000

85 40.000

86 40.O00 84 40.030

1.816 16 1.105 16 5.688 16 --- 16 1.795 19 8.666 19 6.1W 19 S.820 19

97 6oo.000 OB Bo.000 99 I w.000 5.793 e 4 5.838 83 4.169 I 85 .La w.000 .01 60,000 a? I 60.000 .os Iw.000 .04 60.000 a6 MI,000 . D 6 60,000 I

ssol "I 900

PP hr had barn .oormartsd at t h e of lruhllatlon In I1titud*

20 NACA RM E50A31

- 1 P

3 4 6

6 7 8 9

10 11 18 w 14

16 16 17 16

Po 19

81

e4 23

86

86

p8 87

29

50 31 38 55

p a

a4 38 z-6 37 38

39 40 41 48 43 44

46 46 47 48 40

BO

lia 61

03 54

W

6T 66

89 -

w .?e a e.- 7.117 1.788 1.593 1.476

m s.191 1.891 1.893

00

09

2.659 1.K29 1.4U 1.406 1.416

1.686 00

1.491

z.4m 1.410

1.- 1 . S 6 1.076 1.379

1.860 1.809 1.411 1.384

1.C4I 1.794 1 .U6 1.3W 1.576

7 .m 00

6 .W4 1 .Ern 1.420 1.438

1.464 1.4pB 1.599 1.- 1.402

1.666 1.464 1.466 1.411 1.W

t.041 1.846 1.438 1.101

NACA RM E50A31 21

89

9 1 BB

9s 94

96

97

99

m

m

m

101 100

lop

105

106 104

1od

M

108 109 I

60,ooO 1.sa 10,sea q 4 7 s (00 4620 60,000 130 10.743 U . 8 9 4 498 5580 w , m 1.s0 lI,m 12,486 5JM 6oBo

60.000 1.70 10.8e4 U,Qa W 6517 60,OW 1.m 10,lW 10,788 6&8 4 S U 80,OW 1.70 Il.Sp5 1p.279 788 m

I r

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

I 361- /334 I 1253

. .

... .. .. .

a L

. . . . . . . . . . .

N P

I 8 8 S i - I Q F .. . . . . . .

, . . . . . . .

NACA RM E 5 0 A 3 l

Engine speed, N, rpm

Figure 4. - Effect of altitude on jet thrust. Ram pressure ratio, 1.30.

25

26 NACA RM E50A31

3000

2000

1000

0

-1000 4,

Engine speed, N, rpm

Figure 5. - Effect of altitude on net thrust. Ram pressure ratio, 1.30.

.. .

NACA RM E50A31 27

80 Al t i tude

( f t 1.

0 20,000 0 0

70

n 60,000

60

50

40

30

20

10

T 0 4,000 6,000 8,000 10,000 12,000

I Engine. speed, N, rpm

Figure 6. - E f f e c t o f a l t i t u d e on a i r consumption. Ram pressure r a t i o , 1.30.

28

4000

3000

2000

1000

4,000 0 "

NACA RM €=A31

6,000 8,000 10,000 12,000

Engine speed, N, rpm

Figure 7. - Effect of al t i tude on fuel consumption. Ram pressure r a t i o , 1.30.

NAC A RM E50A3 1 29

Engine speed, N, rpm.

Figure 8. - Effect of altitude on net-thrust speol f io fuel consumption. Ram pressure ratio, 1.30.

30 NACA RM E50A31

1800

1600

1400

1200

1000

800 4 ?

E n g i n e speed, Ei, rpm

Figure 9. - Effect of altitude on tail-pipe indicated gaa temperature. Ram preesure r a t i o , 1.30.

NACA RM E 5 0 A 3 1 31

Engine apeed, N, rpm

Figure 10. - Effect. of ram pres su re r a t io on jet thrust. Altitude, 30,000 feet.

32

3000

2500.

2000

3 - 1500. E:

F - 5 .Q 1000 - +,

500 -

-

0-

-

-500 - 4 000

NACA RM E50A31

6 , 000 a,ooo 10,000 12,000 14,000 Engine speed, N, r m

Figure 11. - Effect of ram pressure ratio on net thruet . A l t ituue, 30,000 feet.

NACA RM E 5 0 A 3 1 33,

Engine speed, N, rpm

Figure 12. - Effect of ram pressure ratio on air consumption. Altitude, 30,000 feet.

34 NACA RM E50A3l

4500

4000

3500

3000

2500

2000

1500

1000

500

0 - 4 , 000 6,000 8 , 000 10,000 12,000 14,000

Engine speed, N, r p

Figure 13. - Effect of ram pressure ratio on fuel con- sumption. Altitude, 30,000 feet.

c

E50 A 3 I 35

4,000 6,000 8 000 10,000 12 9 000 14,000

E n g i n e speed, N, rpm

Figme 14, - Effect of ram presaure ratia on net-twust SPeoific fuel ConSUmption. Altitude, 30,000 feet.

36

C r 1800 F

600 t 4,000

Engine apeed, N, rpm

NACA RM E50A31

Figure 15. - ETfect of ram pressure r a t i o on t a i l -p ipe indicated gas temperature. Altitude, 30,000 f ee t .

NACA RM E50.431 37

7000

6000

5000 ft

" * 10

E - 4000 .P

3 m k 5 u - 3000 a 'El a? -P

al 0

k k l2 2000

1000

0 49

Corrected engine speed, N / S , rpm

Figure 16 . - Effect of a l t i t u d e on corrected jet thrust. Ram pressure ratio, 1.30.

38 NACA RM E50A31

4,000 6,000 8,000 10,000 12,000 14,000

Corrected engine speed, N f i , rpm

Figure 17. - Effect of altitude on corrected net thrust. Ram pressure r a t i o , 1.30.

NACA RM E50A31

90

80

70

60

50

40

30 4,000 6,000 8,000 10,000 12,000 14,000

Corrected engine speed, N / G , rpm

39

Figure 18. - Effect of a l t i t ude on corrected air consumption. Ram pressure ra t io , 1.30.

40 NACA RM E50A3f

8000

7000

6000

5000

4000

3000

2000

1000

0 4,

Altitude ( f t )

0 . 0 0 20,000

8 i- 1 00 8,'

Corrected engine speed, N/$, rpm

' & 14,000

Figure 19. - Effect of altitude on corrected fuel consumption. Ram pressure ratio, 1.30.

NACA RM E50A31 41

P rl v h

Corrected engine speed, N/$, rpm

Figure 20. - Effect of altitude on corrected net-thrust speolf ic fuel . consumption. Ram pressure ratio, 1.30.

4 2 NACA RM E50A31

1600

1400

12 00

1000

800

600 4,000 6,000 .8,000 10,000 12,000 14,000

Corrected engine speed, N / G , rpm

Figure 21. - Effect o f a l t i t u d e on cor rec ted t a i l -p ipe indicated gas temperature. Ram pressure r a t i o , 1.30.

NACA RM EWA31

loo0L 4,000 0 Corrected engine speed, N / g , rpm

(a) Corrected j e t thrust, FJ/~.

43

Figure 22. - Effeot of ram pressure ratio on corrected jet thrust. Altitude, 30,000 feet.

44 NACA RM E 9 A 3 1

10 , 000 ~

9 * 000

8 000

-

7 * o m

6,000 -

-

6,000 -

-

4,000 -

-

3,000 4,000 6,000 8,000 10,000 12,000 14,000

Corrected engine speed, IT/*, rpn F : +PnAV

(b) Corrected jet-thrust parameter, J " 6

Figure 22. - Concluded. Effect of ram pressure r a t i o on corrected j e t thrust. Altitude, 30,000 feet.

NACA RM E50A3I 45

Corrected engine speed, N/*, r p m

Figure 23. - Effect of ram pressure r a t i o on corrected net thrust. Altitude, 30,000 feet.

46 NACA RM E a A 3 1

Corrected engine speed, N / G , rpm

F i g w e 24. - Effec t o f ram pressu re r a t io on corrected a i r oonsunption. Altitude, 30,000 fe8t.

NACA Rh4 E5OA3I 47

Corrected engine. speed, N / 6 , rpm

Figure 25. - Effect of ram pressure r a t i o on corrected fuel consumption. Altitude, 30,000 f e e t .

48 NACA RM E50A3 I

"- 4 , 000 6 , 000 8,000 10,000 12,000

Corrected engine speed, N / 6 , rpm

"

14 , 000

Figure 26. - Effeot of ram pressure ratio on corrected net- thrust spec i f ic fue l consumption. Altitude, 30,000 f ee t .

NACA RM E50A31 49

2000

0

E 800 0 V

600 - 4,000 6 , 000 8 000 10 J 000 12 , 000 14 , 000 Corrected engine speed, N/', rpm

Figure 27. - E3feot of ram pressure ratio on corrected ta i l - pipe indicated gas temperature. Alt i tude, 30,000 feet.

50 . . NACA RM E50A3 I

Engine speed, N, rpm

Figure 28. - Effect of' je t -nozz le si20 on tail-pips total pressure. A l t i t u d e , 30,000 feet; ram pressure ratio, 1.70.

. . ."

. . . .

52 NACA RM E50A31

Engine speed, N, rpm

Figure 30. - Effeot of jet-nozzle s i z e on turbine total- pressure ratio. Altitude, 30,000 feet) ram pressure ratio, 1.70.

. NACA RM E50A3I 53

.

.

90

80

70

60

50

40

30

20 4, 000

~

6 , 000 8,000 10,000 12 , 000 14,000 Engine speed, W, rpm

Figure 31. - BPieot of jet-nozzle s ize on enthalpy rise across wmpressor. Altitude, 30,000 feet; ram pressure ratio, 1.70.

54 NACA RM E!jOA31

0 e;

Figure 32. - Effeot of .let-nozzle size on turbine-Inlet total . - a8 temperature. Altitude, 30,000 feet; ram pressure ratio, f.70.

NACA RM E50A31

.

55

Engine speed, N, rpm

Figure 33. - Effect of jet-nozzle size on t a i l - p i p e in- dicated gas temperature. Altitude, 30,000 feet; ram pressure ratio , 1.70.

. 55 NACA RM E50A3 I

4,000 6,000 8,000 10,000 12,000 14,000 Eagine speed, X, rpm

Figure 34. - Effect of jet-nozzle size on afr consumption. Altitude, 30,000 Zeet; ram pressure ratio, 1.70.

. -. .

.

c

NACA RM E50A31 57

4,000 6 , m 8,000 10,000 12,000 14,000

Engine speed, N, rpm

Figure 85. - E3fect of jet-nozzle size on fuel consumption. Altitude, 30,000 feet; ram pressure ratio, 1.70.

30 NACA RM E50A31

Figure 36. - Effect of jet-noeele . s i z e on- jet thruet. Altitude, 30,000 feet ; ram pressure ratio, 1.70.

. NACA RM E50A,3I 59

2500

2000

1500 rl P

a

k

.P 1000

c .r

i .u * ’ 500

0

-500 4,

Engine speed, R , rpm

Figure 37. - Effect of jet-nozzle size on net thrust. Altitude, 30,000 f ee t ; rem pressure ratio, 1.70.

60 NACA RM E50A31

A

Engine speed, N, rpm

Figure 38. - Effect of jet-nozzle- s i z e o n ne t - th r - s t spec i f i c fuel consumption. Alt i tude, 30,000 feet; ram pressure ra t io , 1.70. .

4