I

LEASEDATE U JV1g4 ( ; RESTRICTED COPY No. RM No. E8128b

NACA

RESEARCH MEMORANDUM ALTITUDE -WIND-TUNNEL INVESTIGATION OF WESTINGHOUSE

19B-2, 19B-8, AND 19XB-1 JET-PROPULSION ENGINES

ifi - PERFORMANCE AND WINDIVII L LING DRAG C HARA C TE RIS TICS

By William A. Fleming and Robert 0. Dietz, Jr.

Lewis Flight Propulsion Laboratory Cleveland, Ohio

CLASSIFICATIoN CANCELLED

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F IL C COP NATIONAL ADVISORY COMM1WE&nI

FOR AERONAUTICS Advory

a November 26. 1948 — - !hir,tGfl. ft fl

RESTRICTED

https://ntrs.nasa.gov/search.jsp?R=19930093796 2020-07-04T22:24:24+00:00Z

MACA PM No. EBJ28b &iU 1?EL NAUONAL ADVISORY COMr1u FOR AEROMAUTICS

RESEARCH MEMORANDUM

ALTflJDE -WIND -LUNNJ1 ThV13TIGATION OP WESTINOUSE

19B -2, 19B-8, AND 19Th-i JET-PROPULSION Er.INES

III - PERFORMANCE AND WIT1D1TTJ.TG DRAG CHARACTERISECS

By William A. emi.ng and Robert 0. Dietz, Jr.

SITh'1A.RY

The performance characteristics of the l9B-8 and 19XB-1 turbojet engines arid the windxilling drag characteristics of the 19B-8 engine were determined in the MACA Cleveland altitude wind tunnel. The l9B engine is one of the earliest experimental. Westinghouse axial-flow engines. The 19Th-i engine is an experimental prototype of the Westinghouse 19XB series, having a rated thrust of 1400 pounds. Iiuprovements in performance and. operaticEal characteristics have resulted In the 19XB-2B engine with a rated. thrust of 1600 pounds. The investigations were conducted on the 19B-8 engine at simulated altitudes from 5000 to 25,000 feet with various free-stream rain-pressure . ratios anti on the 19Th-i engine at simuThted altitudes from 5000 to 30,000 feet with approximately static free-stream conditions.

Data for these two enginos are presented to show the effect of altitude, free-stream ram-pressure ratio, and. tail-pipe-nozzle area on engine performance. A 21-percent reduction in tail-pipe-nozzle area of the.19B-8 engine increased the jet thrust 43 percent, the net thrust 72 percent, and the fuel consumption 64 percent. An increase in free-stream ran-pressure ratio raised the et thrust and the air flow and lowered the net thrust throughout the entire range of engine speeds for the 19B-8 engine. At similar operating condi-tions, the corrected. jet thrust and corrected air flow were approxi-mately the same for both engines, and. the corrected specific fuel consumption based on jet thrust was lower for the 19XB-1 engine than for the lOB-S engine.' The thrust and air-flow data obtained with both engines at various altitudes for a given free-stream rain-pressure ratio were generalized to standard sea-level a1nospheric conditions. The performance parameters involving fuel consumption generalized only at high engine speeds at simalated altitudes as high as 15,000 feet. The winthnifling drag of the 19B-8 engine Increased rapidly as the airspeed was increased.

F1LE

2

NACA RM No. E8328b

IN']UCTION

Ixwestitions were conducted in the NACA Cleveland. altitude wind tunnel during October and November, 1944, to determine the performance and operational characteristics of the 19B .-2, 19B-8, and 19KB-i turbojet engines. Operational characteristics of all three engines are presented in reference 1 and the performance char-acteristics of the turbine in the 19B-8 engine are presented in reference 2. Performance characteristics of the 19B-8 and 19KB-i turbojet engines over a range of siniulated flight conditions are presented herein. Data.for the 19B-2 engine are omitted because it is an earlier production model of the 19B-8 engine.

The effects of altitude, free-stream rain-pressure ratio, and tail-pipe-nozzle area on engine performance are shown. The perform.-ance data are generalized to show the applicability of methods used to determine performance at any altitude from rims at a given altitude. Windmiiiing drag characteristicsof the 19B-8 engine are also presented.

Results, are shown for the 19B-8 engine at simulated altitudes from 5000 to 25,000 feet at free-stream ram-pressure ratios up to 1.155 and for the 19KB-i engine at simulated altitudes from 5000 to. 30,000 feet at free-stream ram-pressure ratios below.l.Ol.

DESCRIPTION OF •ENGIS

19B-8 engine. -. The l9B-8 turboje.t engine has a static sea-level rating of 1365 pounds of. thrust at an engine speed of 17,500 xpm. At this rating the engine has an air flow of 28 pounds per second and a fuel consumption of approximately 1800 pounds per

hour. The over-all length of the engine is l04- inches with the

adjustable tail cone extended; the maximum dianieter is 20 inches and the total weight is 825 pounds. 'The maximum diameter does not include the accessory oup mounted on the front bearing support at the compressor inlet.

This engine has a six-stage axial-flow compressor, an annular-type combustion chamber, and a sing1e-stae turbine. -kie1 is injected downstream into the combustion chaniber through 24 nozzles mounted circumrerentially in the manifold at the fO±ward end of the combustion chairiber. The gases leaving the turbine pass through a

NACA RN No. E8328b 3

tail pipe equipped with a movable inner cone, ,hich has a total axial travel of 5 inches. When the inner cone is moved from the "i" or for ard position to the tt4_inchesouttt position, the tail-pipe-nozzle area.is decreased from 135 to 106 square inches. The change in area is nearly linear with respect to the travel of the tail cone.

In order to cool the oil for lubricating the engine bearings and accessory drive, a cylindrical oil cooler 2 feet long with an inside diameter of 14 inches is attached to the front of the engine. The inner wall of the cylinder provides the heat-exchanger surface for the oil.

19XB-1 engine. - The l9Th-1 turbojet engine is a modification of the 19B-8 engine that contains a ten-stage axial-flow compressor and has the same over-all engine dimensions. The tail pipe has a fixed inner cone and the tail-pipe-nozzle area is 103 square inches. The 19XB-1 engine has a sea-level rating of approximately 1400 pounds static thrust at a rated engine speed of 16,500 rpm. At this rating the air flow is 29 pounds per second and the fuel consumption is approximately 1600 pounds per hour.

ThETAILATION AND PR0C)UPE

Each engine was installed in a nacelle mounted under a stub wing that extended from one side of the 20-foot-diameter test sec-tion of -the wind tunnel. (See fig. 1.) The engine air was admitted at the front of the nacelle, and a small duct beneath the J.eading edge of the wing admitted air to cool the outer wall of the combus-tion chamber and the tail pipe. For the runs at approximately static. conditions, when the free-stream ram-pressure ratio was below 1.01, the part of the cowling enclosing the tail pipe and the rear of the combustion chamber was removed because the cooling-air flow through the cow1ing.as inadequate.

Survey rakes were installed at several stations in the engine (fig. 2) in order to obtain temperature and pressure measurements from which the component and over-all performance of the engine could be determined. The methods used to calculate performance parameters are presented in the appendix.

Throughoutthe part of the investigation discussed, a 6-mesh screen of 0.032-inch-diameter steel wire was attached to the front flange of the oil cooler to protect the compressor from flying objects in the tunnel.

4 NACA RN No. E8328b

The 19B-8 engine was operated at simulated altitudes from 5000 to 25,000 feet with free-stream ram-pressure ratios up to 1.155; and. the 19XB-1 engine, from 5000 to 30,000 feet at approximutely static free-stream conditions. The tunnel temperature was held at approxi-mutely NACA standard values for each flight condition. At each simulated flight condition, the engine was run over the full range of operable engine speeds. The fuel specified for these engines, 62-octane unleaded gasoline, was used.

Dtring the investigation, the combustion chambers were modified several times (see reference 1) in an attempt to increase the operating range of the engines at high altitudes. Because none of the modifications made any siiificant improvement in the operating range of the engine, all the data presented are for the engines with the standard combustion chambers as received from the manufacturer.

SYMB0L

The, following syn1bol are used in this paper:

A cross-sectional area, square feet

B thrust-scale reading, pounds

installation drag coefficient

Cp specific heat of gas at constant pressure, Btu per pound °R

windmilling drag, pounds

Fj jet thrust, pounds

F net thrust, pounds

f/a fuel-air ratio

g acceleration of gravity, 32.2 feet per second per second

J mechanical equivalent of heat, 778 foot-pounds per Btu

m mass flow, slugs per second

N engine speed, rpm

WCA BM No. E8J28b 5

P total pressure, pounds per square foot absolute

p static pressure, pounds per square foot absolute

R gas constant, 53.4 foot-pounds per pound °R

S wing area, square feet

T total temperature, °B

T indicated. teniperature, °R

t static temperature, °R

thp net thrust horsepower

V velocity, feet per second

velocity, miles per hour

Wa air flow, pounds per second

Wf fuel consumption, pounds per hour

Wf/Fj specific fuel consumption based on jet thrust, pounds per hour pound thrust

Wf/Fn specific fuel consumption based on net thrust, pounds per hour pound thrust

Wf/thp specific fuel consumption based on net thrust horse-power, pounds per hour per thrust horsepower

ratio of specific heats for gases

S ratio of absolute tunnel static pressure to absolute static pressure of NPLCA standard atmospheric con-ditions at sea level, p0/2116

e ratio of absolute tunnel static temperature to absolute static temperature of NACA standard atmospheric conditions at sea level, t0/519

p mass density, slugs per cubic foot

6

NACA RM No. E8J28b

Subscripts:

a air

f fuel

gas

j station at which static pressu.re in jet reaches ambient static pressure

0 ambient free stream

1 cowl inlet

2 compressor inlet

5 turbine outlet

The following performance parameters generalized to NACA standard atmospheric conditions at sea level were used:

D /8 corrected wind.illing drag, pounds

Fj/8 corrected jet thrust, pounds

F/8 corrected net thrust, pounds

(f/a)/O corrected. fuel-air ratio

N/ \/ corrected engine speed, rpm

corrected true airspeed, miles per hour

Wa corrected air flow, pound.s per second

Wf/(8 /) corrected fuel consumption, pounds per hour

Wf/(Fj'\[ë) corrected specific fuel consumption based on jet thrust, pounds pel' hour per pound jhrust

corrected specific fuel consumption based on net thrust, pounds per hour per pound thrust

Wf/thp corrected specific fuel consuiimtjon based on net thrust horsepower, pounds Der hour per thrust horsepower

NACA R No. E8J28b

7

PFORMA.NCE CHRACTEIRISTICS

Effect of Inlet Screen

The screen Installed at the cowl Inlet for the wlnd-ttumel investigation caused a total-pressure loss between the free stream and the compressor inlet that varied with the corrected engine speed. The ram-pressure ratios used. in this report are based on measurements of free-stream total pressure instead of compressor-inlet total pressure. Because the tunnel-test sectionairspeed was used to calculate net thrust and net thrust horsepower, the engine-perfornnce characteristics involving these parameters were slightly affected. The relation between corrected engine speed. and the ratio Of compressor-inlet total pressure to free-stream total pressure is shown in figure 3. The curve shown In this figure was determined from data obtained throughout the entire range of flight conditions Investigated. The compressor-inlet ram-pressure ratio at any engine speed ny be obtained for the data presented by zwfltiplying the free -stream ram-pressure ratio by the value of P2/P0 obtained from figure 3 for the corresponding corrected engine speed.

Effect of Altitude

1OB-8 engine. - The 19B-8 engine perfornce at sinn.aated altitudes of 5000, 15,000 and 25,000 feet, a free-stream rein-pressure ratio of 1.013, and with a tail-pipe-nozzle area of 135 square inches, is shown in fIgure 4. Because of the high mm-inrum speed of the engine at altitudes above 17,000 feet (refer-ence 1), no data could be shown for an engine speed below 15,500 rpm at an altitude of 25,000 feet. The variations In jet thrust, not thrust, fuel consumption, air flow, and specific fuel consumption based on net thrust with increasing altitude are shown In figures 4(a.) and 4(e), respectIvely.

Above an engine speed of 13,000 rpm the fuel-air ratios at simulated altitudes of 5000 and 15,000 feet were the same (fig. 4(f)). At an altitude of 25,000 feet, howevor, the fuel-air ratio through-out the entire operable range of the engine was higher than at 5000 and 15,000 feet. The high fuel-air ratio obtained at 25,000 feet was probably related to the reduction in combustion efficiency. A low combustion efficiency at a simulated altitude of 25,000 feet was indicated by the narrow operating range of the engine as limited by combustion blow-out. At engine speeds below

8 NACA RM No. E8J28b

13,000 rpm, the iücrease in fuel-air ratio at en altitude of 15,000 feet above that at 5000 feet was also caused by the reduction in combustion efficiency that accompanied an increase in altitude.

19XB-]. engine. - The l9-1 engine perfornE.nce at appror1n.tely static conditions (free-stream rain-pressure ratio less than 1.01) and. siTmilAted. altitudes of 5000 to 30,000 feet is shown In figure 5. WT4imnn engine speed was limited, by high tail-pipe temperatures and. above an altitude of 15,000 feet the Tninlrmnn engine speed ben to increase. The reduction in jet thrust, fuel consumption, and. air flow that resulted from an increase In altitude, Ia shown in figures 5(a) to 5(c), respectively. No net-thrust curve is presented. inasmuch as these data were obtained, at app orlite1y static cond.I - tions where the jet and net thrusts are equal. Variation of specific fuel consumption based on jet thrust with altitude and engine speed Is shown in figure 5(d). For high engine speeds,, the specific fuel consumption was slightly higher at altitudes of 25,000 and 30,000 feet than at the lower altitudes. The effect of altitude on the relation between engine speed and. fuel-air ratio is shown in figure 5(e).

Effect of Rain-Pressure Ratio

The effect of varying the free-stream ram-pressure ratio from 1.012 to 1.155 throughout the operable speed range of the 19B-8 engine with a tail-pipe-nozzle area of 106 square inches at an altitude of 15,000 feet is presented in figure 6. The maximum

-engine speed obtainable with the tail cone 4 inches out (tail-pipe-nozzle area, 106 sq in.) was limited to about 16,500 rpm by high •tail-pipe temperatures. Increasing the ram-pressure ratio raised the jet thrust (fig. 6(a)) and. the air flow (fIg. 6(c)) and lowered the net thrust (fIg. 6(b)) throughout the entire range of engine speeds. At all engine speeds below 16,500 rpm the fuel consumption was lowered as the mn-pressure ratio was increased (fig. 6(d)).

At low engine speeds, the specific net thrust Increased, as the ram-pressure to 1.155 (fig. 6(e)). Above a speed of specific fuel consumption increased only pressure ratio from 1.012 to, 1.122. The based on net thrust horsepower decreased pressure ratio throughout the entire ran (rig. 6(f)). At an engine speed of 16,5

fuel consumption based on ratio was raised from 1.012 14,000 rpm, however, the for a change in rain-specific fuel consumption with Increasing ran-ge of engine speeds 00 rpm, the specific fuel

NACA RN No. E8328b

9

consumption based on net thrust increased from 1.51 at a ram .-pressure ratio of 1.012 to 1.62 at a ram-pressure ratio of 1.155 (flg. 6(e)) and the specific fuel consumption based on net thrust horsepower decreased from 5.25 to 1.85 for a corresponding increase in ram-pressure ratio (fig. 6(f)). The relation between specific fuel consumption based on net thrust and net thrust Is shown in fIgure 6(g).

The fuel-air ratio (fig. 6(h)) was reduced throughout the entire range of engine operation as the ram-pressure ratio was increased.

Effect of I1-PIpe-Nozzle Area on PerfornE.nce

The effect of a 21 .-percent reduction in tail-pipe-nozzle area (from 135 to 106 sq. in.), which was obtained by nioving the tail cone from the in position to the 4-inches-out position, on the performance of the 19B-8 engine at a free-stream ram-pressure ratio of 1.131 and at a simflated altitude of 10,000 feet, Is shown in fIgure 7. With a tail-pipe-nozzle area of 106 square inches, the maxinm engine speed was limited to 16,500 rpm by the turbine-inlet temperature (1500° F). The reduction in tail-pipe-nozzle area increased the jet thrust (fig. 7(a)), net thrust (fig. 7(b)), and fuel consumption (fig. 7(c)) and lowered the air flow (fig. 7(d)) throughout the range of engine speeds. At an engine speed of 16,500 rpm, the reduction in nozzle area increased the jet thrust 43 percent (fig. 7(a)), the net thrust 72 percent (fig. 7(b)), and the fuel consumption 64 percent (fig. 7(c)).

Throughout the range of engine speeds, the specific fuel con-sumption based on net thrust (fig. 7(e)) and the specific fuel consumption based on net thrust horsepower (fig. 7(f)) were lower with a tail-pipe-nozzle area of 106 square Inches than with an area of 135 square inches. At an engine speed of 16,500 rpm, the specific fuel consumption based on net thrust was approximately 6 percent higher with a tail-pipe-nozzle area of 135 square inches than with an area of 106 square inches. Changing the tail-pipe-nozzle area had no effect on the relation between the specific fuel consumption based on net thrust and the net thrust (fig. 7(g)).

The reduction In tail-pipe-nozzle area increased the back pressure on the turbine and as a result the engine required a higher turbine-inlet temperature, that Is, a higher fuel-air ratio to main-tain a given engine speed. The increase in fuel-air ratio resulting from a reduction in tail-pipe-nozzle area is shown in fIgure 7(h).

10

NACA RN No. E8J28b

The concept of flow similarity and the application of dimen-sional analysis to the performance of turbojet engines has led, to the development of the pressure and temperature reduction factors 8 and. 0 with which data obtained at several altitudes may be gener-alized. Data thus far presented that were obtained at various alti-tud.es and free-stream ram-pressure ratios are generalized, to NACA standard conditions at sea level in order to test the applicability of the factors 8 and U to the performance of the l9B-8 and 19KB-1 turbojet engines.

Application of the factors 8 and 0 give the following gen-eralized performance variables: corrected engine speed, N/4; corrected jet thrust, F 1/6; corrected net thrust, FJ6; cor-rected air flow, (Wa/)/8, corrected fuel consumption, wf/(8/); corrected specific fuel consumption based on jet thrust, Wf/(Fjy'); corrected specific fuel consumption based on net thrust, Wf/(F); corrected specific fuel consumption based on net thrust horsepower, w/thp; and corrected fuel-air ratio, (f/a)/0.

The generalized performance variables that contain fuel con-sumption exclude the effects of variations in combustion efficiency, which is a component of the factor of thermal expansion of the working fluid included in the dimensional analysis. This exclusion adnittedly lessens the possibilities of successfully generalizing the data in cases where the corrlbustion efficiency changes.

The effect of altitude on the generalized performance of the 19B-8 and 19KB-i engines is shown in figures 8 and 9, respectively. The corrected jet thrust (figs. 8(a) and 9(a)), the corrected net thrust (fig. 8(b)), and the corrected air flow (figs. 8(c) and 9(b)) generalized to a single curve for each engine. The parameters involving fuel consumption, that is, corrected fuel consumption (figs. 8(d) and 9(c)), corrected specific fuel consumption (figs. 8(o), 9(d), and. 9(e)), and. corrected fuel-air ratio (figs. 8(g) and 9(f)), generalized only in the upper range of engine speeds for simulated, altitudes as high as 15,000 feet.

Performance data obtained at various ram-pressure ratios did not generalize to a single curve inasmuch as the factors 8 and ê are not meant to correct for the changesin cycle efficiency that accompany variations In ram-pressure ratios. The effect of ram-pressure ratio on the generalized performance of the l9B-8 engine

NACA RM No. E8J28b

II'

is shown in figure 10. The perfornmnce paraieters involving cor-rooted jet thrust (fig. 10(a)), corrected net thrust (fig. 10(b)), corrected air flow (fig. 10(c)), corrected fuel consumption (fig. 10(d)), corrected specific fuel consumption based on net thrust (figs. 10(e) and. 10(r)), specific fuel consumption based on net thrust horsepower (fig. 10(g)), and corrected fuel-air ratio (fig. 10(h)) produced a family of curves for various free-stream ram-pressure ratios.

In the use of generalized perfornnce curves, the effect of ram-pressure ratio on engine perfoxnce at various altitudes can be determined only by use of the entire family of curves. It is impossible to predict the perfornnce parameters involving fuel consumption at low engine speeds or at hi&i altitudes because the effects of varying conthustion efficiency are excluded front the factors used in this report.

A comparison of generalized perfornntnce data of the 19B-8 and 19B-1 engines at a corrected engine speed of 17,000 rpm is pre-sented in the following table:

19B-8 I 19XB-1

il-pipe-nozz1e area, sq in. Corrected engine speed, rpm Corrected jet thrust, F j/8, lb Corrected air flow, Waifl/5, lb/sec Corrected fuel consumption, W/8/ lb/hr Corrected specific fuel consumption based on

jet thrust, Wf/(FjI\/), lb/(hr)(lb thrust) Corrected fuel-air ratio, (f/a)/e

106 103 17,000 17,000

1370 1375

27.0 26.5

1965 1825 1.435 1.328

0.0202 I 0.0192 Data presented in this table were obtained from figure 9 for

the 19JCB-1 engine at static conditions and from figure 10 for the 19B-8 engine at a free-stream ram-pressure ratio of 1.012. At static conditions, the ejector effect of the jet produced a velocity in the tunnel test section at a n.xixm.un engine speed equivalent to a free-stream ram-pressure ratio of approxitely 1.012. For the conditions given in the table, the compressor-inlet ram-pressure ratio for both ongines was 0.938. Generalized data are used because the inlet-air, temperatures were different for the two engines.

12 NASA RM No. E8328b

The corrected. jet thrust and the corrected air flow of both engines at a corrected engine speed of 17,000 rpm were approxinE.tely the same. The corrected specific fuel consumption based on jet thrust was lower for the 19Th-i engine than for the 19B-8 engine.

The n itude of the wind.mifling drag obtained when an engine is inoperative and is allowed to windmill during flight is equal to the change of inomentuni of the air as it passes through the engine. A limited amount of windmnilling drag data was obtained for the 19B-8 engine at simulated altitudes from 5000 to 30,000 feet and tunnel-test-section airspeeds from 176 to 373 miles per hour with the screen installed at the cowl inlet. Generalized values of wind-milling drag (fig. 11(a)) show that the windmnilling drsg Increased rapidly as the corrected true airspeed was increased. A single curve was drawn through the generalized values obtained at various altitudes. The relation between the ratio of windmilling drag tO the iEximum net thrust obtainable with a tail-pipe-nozzle area of 135 square inches arid the tunnel-test-section true airspeed Is shown in figure 11(b). These data show that at any altitude the wind-milling drag of the engine at an airspeed of 200 miles per hour was 2.5 percent of the maximum net thrust obtainable at this airspeed and. that at an airspeed of 375 miles per hour the windinilling drag was U percent of the me.xlmumn net thrust obtainable at this air-speed. The generalized values of air flow obtained with the engine wim3inifling (fig. 11(c)) fell on a single curve for all altitudes investigated.

StIMMARY OF RL

om an investigation in the Cleveland altitude wind tunnel of the perfornance characteristics of the l9B-8 and 19Th-i turbojet engines and the windmifling drag characteristics of the 19B-8 engine, the following results were obtained:

1. A reduction in tail-pipe-nozzle area of the 19B-8 engine from 135 to 106 square inches at a simulated altitude of 10,000 feet and a free-stream ram-pressure ratio of 1.131 gave an Increase in jet thrust, net thrust, and fuel consumption with a reduction in air flow and specific fuel consumption based on net thrust throughout the entire range of engine speeds. This reduction in tall-pipe-nozzle area at an engine speed of 16,500 rpm raised the jet thrust 43 percent, the net thrust 72 percent, and the fuel consumption 64 percent. -

NACA RM No • E8J2 Sb 13

2. With a tail-pipe-nozzle area of 106 square inches on the 19B-8 engine, an increase in free-stream ram-pressure ratio from 1.012 to 1.155 at a simulated altitude of 15,000 feet raised the jet thrust and. air flow and lowered the net thrust throuiout the entire range of engine speeds. The fuel consumption was lowered at all engine speeds below 16,500 rpm.

3. At similar operating conditions, the corrected jet thrust and the corrected air flow were .approxlte1y the same for both engines and the corrected specific fiie1 consumption based on jet thrust was lower for the 19Th-1 engine than for the 19B-8 engine.

4. With the use of generalizing pressure and temperature factors jet-thrust, net-thrust, and air-flow data obtained from both the l9B-8 and 19)-1 engines at a given ram-pressure ratio and tail-pipe-nozzle area, and at any altitude could be used to estinte the perfornmnce of the engines at any other altitude. The perform-ance parameters Involving fuel consumption generalized only at hli engine speeds at simulated altitudes up to 15,000 feet. Above this altItude these parameters did not generalize at any engine speed.

5. The wlnd.mlfling drag of the l9B-8 engine while it was inoperative amounted to 2.5 percent of the nximum net thrust of the engine at an airspeed of 200 miles per hour, and. 11 percent of the me.ximum net thrust of the engine at an airspeed of 375 miles per hour.

Lewis Fliit Propulsion laboratory, National Advisory Coimnittee for Aeronautics,

Cleveland, Ohio.

NACA PM No. E8J28b

APPEfl - METEODS OF CALCULATION

Temperature. - A sample thermocouple of the type used was call - brated. cold. up to a ch number of 0.8. This calibration showed the thermocouple measured. the static tentperatu.re plus approximately 85 percent of the adiabatic temperature rise due to the impact of the air on the thermocouple. Static temperature may be 'determnlned from indicated temperature by applying this factor to the adiabatic relation between temperature nd pressure:

t=

1+O.85[()l]

and the total temperature is

Z fP'\ 7

Tj)

l+0.85I\ 7 _l] L 1

In order to determine the static temperature of the jet, the assumption was made that no loss occurred in total temperature or total pressure between station 5 in the tall pipe and station j in the jet. When this assumption is made, the static temperature in the jet is

T5 ti =

(P5\\ i

'PO)

The tunnel-test-section static temperature was determined from the indicated temperature measured in the large section of the tunnel just ahead of the test section. Because the velocity in the large section of the tunnel was very low, the indicated temperature measured in the large section is equal to the test-section total temperature. The static temperature in the test section then is

NACA PM No. E8J28b

15

T0 to= 7ol

I •'0

"PC

Tunnel-test-section velocity. - The tunnel-test-section velocity was obtained from

70_i

V0=2Jgct0 ()

-1

Air flow. - The air flaw through the engine was determined from pressure and temperature measurements obtained with four survey rakes at the cowl inlet (station 1, fig. 1). The equation used. to calculate the air flow is

- 71_i

p A1 2J g cp IP1\ 7i Wa = p1 A1 V1 g = i tl •\Pl)

Thrust and. winthnifling drag. - Two methods were used. to cal-culate the jet thrust. For the tests at approxiiinteiy static conditions when the tunnel-test-section velocity was low, the thrust was measured directly on the balance scales. The external drag of the nacelle and. the initial momentum of the inlet air were added to the scale thrust reading, which gave the foflowing equa-tIon for et thrust:

Fj =B + l/2 PO VO2OTj S + ma Vo (1)

Inasmuch as the scale-thrust term B of equation (1) was the njor part of the total thrust and could be measured. within ±2 pounds, the jet thrust could be determined accurately at low test-section velocities. Although variations in the drag coeffi-cient accompanying a change in inlet-velocity ratios could not be determined, an error of as much as 50 percent in the drag coeffi-cient would result in less than a 1-percent error in thrust at test-section airspeeds below 100 feet per second.

16

NCA RM No. E8J2 Sb

For test-section aIrspeeds of 100 miles per hour or more, the jet thrust was determined by using the mass air flow measured at the cowl Inlet, the fuel consumption, and. the jet velocity at the vena contracta In the following equation:

Fj = (ma +mf) V (2)

Calculating the jet thrust with these two methods for the static tests gave a calibration factor that could be applied to equation (2). The value of the calibration factor, which is the ratio of measured scale thrust to calculated tail-pipe thrust, was found to be 1.015. By expansion of equation (2) aM application of this factor to It, the following expression used to calculate the jet thrust for an of the tests except the static tests was obtained:

r -'

Fj = 1.015 mg 2J g cp,j tj - 1]

The net thrust was then determined by subtracting tho free-stream momentum of the inlet air for the jet thrust:

= Fj - ma

The net thrust horsepower was determined from the net thrust and the true airspeed by the expression

1? •r

..n O thp = 550

The wind.millIng drag of the engine was determined from the product of the mass air flow aM the change in velocity of the air that passes through the engine by use of the equation

D = ma (V0 - Vj)

NACA RM No. EBJ28b

17

REVEPENCES

1. Fleming, Williarii A.: Altitude-Wind-Tunnel Investigation of Westinghouse l9B-2, l9B-8, and. 19)CB-1 Jet-Propulsion Engines. I - Operational Characteristics. NACA RN No. E8J28, 1948.

2. Krebs, Richard P., and. Suozzi, Frank L.: Altitude-Wind-Tunnel Investigation of Westinghouse 19B-2, 19B-8, and 19)CB-1 Jet-Propulsion gines. II - Analysis of Turbine Performance of 19B-8 &igino. NACA RN No. E8J28a, 1948.

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NACA RM No. ESJ28b

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0 A.

A.

a

4) 4)

U U)

4' 4)

.1;'

0 .9 U

U, U) 0

U

0 .9

4

I,

4)

1

9: a)

'-4

4' 0

.90 L-7,000

.9

NACA RM No. E8J28b

9.000 11,000 13,000 15,000 17,000 19,000 Corrected engine speed, N/'[, rpm

Figure 3.- Relation between total-pressure ratio across cowl-inlet screen and corrected engine speed for range of simulated altitudes and flight speeds.

100C

'-4

... 60C

4'

•' 40C 4) a'

C

w

____ ____ ____Simulated alt itude

(ft)

0 5,000

o 15,000 25,000

____ ____

___—/

_ V ___

--__ ___ _

00 8000 10.000 12.000 14.000 16..000 l'R.0

NACA RM No. E8J28b

23

Engine speed, N, rpm

(a) Relation between jet thrust and engine speed.

Figure 4.— Effect of altitude on perfornance characteristics of 198.-B turbojet engine. Tall—pipe—nozzle area, 135 square tnches; free—stream ram—pressure ratio, 1.013.

24

NACA RM No. E8J28b

p '.4

a 4,

I

a., a. , &gine speed, N, rpm

(b) Relation between net thrust and engine speed. Ptgure 4. — Continued. Effect of altitude on performance character-

istics of 198-8 turbojet engine. Tail —pipe—nozzle area, 135 square inches; tree—stream ram—pressure ratio, 1.013.

I

NACA RM No. E8J28b

25

1400

1200

1000

'-4

800

0 '-4

t600

r4

400

200

0 6,

H Simulated

_____ altitude _____ ___ - (ft)

0 5,000 o 15,000. O 25,000

_ _ _ —/-_ _ -7--

/

_-__ _ _

' 5 8 10.000 12.000 14 T6 le. )0O Engine speed, N, rpm

(c) Relation between fuel consumption and engine speed. Figure 4.- Continued. Effect of altitude on performance character-

istics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches; free-stream ram-pressure ratio, 1.013.

____ ____ ____ ____ _____ ____Simulated altitude

(ft)

o 5,000

o 15,000 ' 25,000

____ ____

0

__ __ V ___

__ V __7

_

25

20

C.)

U)

.

15 ('I

C '-4 '10 $4

5

0 _____ _____ _____ _____ _____ _____ _____ _____ _____ 6,_._

Engine speed, N, rpm (d) Relation between air flow and engine speed. Pigure 4,- Continued. Effect of altitude on performance

characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches; free-stream ram-pressure ratio, 1.013.

R.00O rrw 18,000

26

NACA RM No. E8J28b

•II1_____ •11____ __

__"I •11--

__

__ • Li___

UI-lip

I III __

It:W..

5.

4, (I)

4) 4• 0

'-4

0 4•

3. a

4) (I)

4, 4) 0)

0

V '1) In

•0

0 '-I

I

3.

2.

0

1.

1.01 I I I I I I I I I 6,000 8,000 10,000 12,000 14,000 16,000

Engine speed, N, rpm

(e) Relation between specific fuel consumption based on net thrust and engine speed.

Pigure 4.- Continued. Effect of altitude on performance characteristics of 198-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches; free-stream rain-pressure ratio, 1.013.

NACA RM No. E8J28b

27

14

0 '.4 4)

1.

• 0l

28

NACA RM E8J28b

• OlE

• 01€

. 01

.010 ____ ____ ____ ____ ____ ____ ____ 6,uvv 10,000 1,U0U 14,000 16,000 18,000

Engine speed, N, rpm

(f) Relation between fuel-air ratio and engine speed.

Figure 4.- Concluded. Effect of altitude on performance character-istics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches; free-stream ram-pressure ratio, 1.013.

Simulated altitude

(ft) 0 5,000 0 15,000

25,000

iiiii

/P

NACA RM No. E8J28b 29

Simulated altitude

(ft)

0 5,000 o 10,000

1200

o 15,000 25,000

V 30,000

I . I

FX

800 0

4600

4) 0)

400

01 I I I I I I I

6,000 8,000 10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

(a) Relation between jet thrust and engine speed.

Figure 5.- Effect of altitude on performance characteristics of 19Th-i turbojet engine. Approximately static conditions.

1

1

•0 r4

0

0.

U

1

30

NACA RM Mo. E8J28b

6,000 8,000 10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

(b) Relation between fuel consumption and engine speed.

Figure 5.- COntinued. Effect of altitude on performance character-istics of 19XB-1 turbojet engine. ApproxImately static conditions.

U

•0 '-4

a

0 r•'J 1C '4

'-4 'C

8,000 10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

25

2C

w

____

S I mu la ted alt .tude

(ft)

0 5,000 o io,000

15,000 25,000

V 30,000

_ ___ - _

-.'- '.-_ø__--

_ _ _

NACA RM No. E8J28b

3'

(c) Relation between air flow and engine speed.

Figure 5,— Continued. Effect or altitude on performance character-istics of 19X8—1 turbojet engine. Approximately static conditions.

32 N/ALA FM No E8J28b

3..

2

1.

1.6,000 8,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm

(d) Relation between specific fuel consumption based on jet thrust and engine speed.

Figure 5. — Continued. Effect of altitude on performance characteristics of 19X8—1 turbojet engine. Approximately static conditions.

4-)

4)

'-4

'-4

r.

4-,

4-)

4)

4-,

'V

0

'V (I)

C 0 4)

I

Simulated altitude

____ ____ ____ ____ (ft) ____ ____

o 5,000 o 10,000

15,000 25,000

V 30,000

0

_____ __-

C____ ____ ____ ____ ____ ____ ____ ____ ____ ____

NACA RM No. E8J28b

33

.016

'4

0

.014

w

ra. .012

.010 6,000

.018

___ ___ ___

Simulated altitude

(ft)

o 5,000 o 10,000

15,000 25,000

V 30,000

___ ___

-__ __

__Iii _8,000

(e) Relation between

Pigure 5. — Concluded. teristics of 19XB—1 conditions.

10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

fuel—air ratio and engine speed.

Effect of altitude on performance charac-turbojet engine. ApproxImately static

34

NACA RM No. E8J28b

Fr e e- Stream ram -

pressure ratio

o 1.012 o 1.122 G 1.155

___ 7 . ____ _

1000

800

r4

600 a-)

a 4)

400

4)

200

08,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm (a) Relation between jet thrust and engine speed. Figure 6.- Effect of free-stream ram-pressure ratio on per-

formance characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet.

NACA RM No. ESJ28b

35

800

.0 '-4

600

4) (n

20C

•1_______________________________________________________________________________________________

- . I- -- - -

•1

_ __ ___I

•_ _____

•__ _________

8,000 10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

Cc) Relation between air flow and engine speed.

Figure 6.— Continued. Effect of free —stream rain—presstu'e ratio on perrormance characteristics of 19B-8 turbojet engine. Tail—pipe—nozzle area, 106 square inches; simulated altitude, 15,000 feet.

0

WA ______

• *i4 hi___ __________ ___

C, a,

0 '-4

'-4 4

• _

Free-s tream ram-

pressure - ratio o 1.012 o 1.122 O 1.155

____ ____ ____ ____ ____ ____ ____ ____ ____

H ________ __

77

) ___ _________ ______ ______ ______ ______

120(

100(

I80(

'-I

a

0 4)

60 0.

401

20

36

NACA RM No. E8J280

8,000 10,000 1Z,000 14,000 16,000 18,000 Engine speed, N, rpm

(d) Relation between fuel consumption and engine speed. Figure 6.- Continued. Effect of free-stream ram-pressure ratio

on perrormance characteristics of 19B-8 turbojet-engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet.

NACA PM N. E8J28b

37

_________________F

I''____ .iI_____

_

_LUI it.________

______

BLI ____

--'-i _

43

U)

43

0

'.4 I '3, 43

43 w

0

V U,

Cd

0 q.4

4.0

3.5

3.0

2.5

2.0

1.5

1.01 I I I I I I I I .1 I I_

8,000 10,000 12,000 14,000 16,000 18,000 &gine speed, N, rpm

(e) Relation between specific fuel. consumption based on net thrust and engine speed.

FIgure 6.- Continued. Effect of free-strem ram-pressure ratio on performance characteristics of 19B-8 turbojet engine. Tai1-pipe-noz1e area, 106 square inches simulated altitude, 15,000 feet.

38

9.(

.o 8.( 1-4

4)

1•

0

(0

0

4) U,

4)

4) Q)

0 .ts

0

3.0

2C

1.0

NACA RM No. E8J28b

__ __ I. __ __ __ ____ ____ ____

__ __ ________

Free—stream ram—

pressure ratio

0 1.012 0 1.122 'Z 1.155

__

____\\

\\

H

6.0

S.0

4.0

8,000 10,000 12,000 14,000. 16,000 18,000 Engine speed, N, rpm

(f) Relation between specific fuel consumption based on net thrust horsepower and engine speed.

Figure 6. — Continued., Effect of free —stream ram—pressure 'ratio on performance characteristics of 19B-8 turbojet engine. Tail—pipe—nozzle area, 106 square inches; simulated altitude, 15,000 feet.

NACA RM No. E8J28b

39

Free-stream ram-

pres sure ratio

o 1.012 o 1.122

1.155

____ ____ ____ ____ ____

____ ____

J

- J-- — 3 I

4.)

•0 —1

'Z 4.0

rl

a

3.

4.) U,

4.)

3.c

C

C 0

2.

0

4.)

0.

2.(

C) 1

—4 C)

Ui

. sfl.# a ,....nj ,.s.., S.,

Net thrust, F, lb

(g) Relation between specific fuel consuiiption based on net thrust and net thrust,

Figure 6.- Continued. Effect of free-stream ram-pressure ratio on performance characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simUlated altitude, 15,000 feet.

.020

.018

.016

.014

$ 0 WI 4)

WI

V r4 w

.012

.010

40

NACA RM No. E8J28b

.008 ____ ____ ____ ____ ____ ____ ____ ____ ____ 8,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm (h) Relation between fuel—air ratio and engine speed.

Figure 6, — Concluded. Effect of free—stream rain—pressure ratio on performance characteristics of 198-8 turbojet engine. Tail—pipe—nozzle area, 106 square inches; simulated altitude, 15,000 feet.

Free—stream ram -

pressure ratio

0 1.012 o 1.122 0 1.155

______ ______ ___

•

liii• _____

/ ___ _ _ _

NACA RM No. E8J28b

41

0 ____ _____ ____ ____ ____ _____ ____ ____ ____ _____ 8,000 10,000 12,)00 14,000 16,000. 18,000

Engine speed, N, rpm

(a) Relation between jet thrust and engine speed.

Figure 7.- Effect of tail-pipe-nozzle area on performance characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, l.11; simulated altitude, 10,000 feet.

1200

1000

800

'-4

a

600

4)

400

20C

_____ _____ _____Tail-pipe-nozzle area (sq in.)

o 155 0. 106

_____ _____ _____ _____ _____

__ __ __p0 __ __

•. 7/ _

_ 1/ _ _ _/7 _ _ _ _

7/7_

__/7 _ _ _

42

NACA RM No. E8J28b

Ta 11-pipe-____ ____ ____ ____ ____ ____ ____ nozzle ____ ____

- area (sq In.)

o 135

o 106

-7

t ______// __

'7-C 8,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm

(b) Relation between net thrust and engine speed. Figure 7.- ContInued. Effect of tail-pipe-nozzle area on per-

formance characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude, 10,000 feet.

1000

800

.0 r1

600

4) U)

400 43 0) z

200

NACA RM No. E8J28b

43

1400

1200

1000

-.4

' 800

0 -I 4-) p.

600

Tail-pipe-

____ ____ • ____ ____ ____ ____ ____ I nozzle

I area

• ( (sq In.)

0 135

0 106

/•

_____//____

-.4

400

8,000 10,000 12,000 14,000 16,000 18,000 Engine speed, N, rpm

Cc) Relation between fuel consumption and engine speed. Figure 7.- Continued. Effect of tail-pipe—nozzle area on per-

formance characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude, 10,000 feet.

44 NACA RM No. E8J28b

0 41) 0

a

3

0 ,-'4

-4

2

2(

____ ____Tall-pipe-nozzle area (sq In.)

o 135 o 106

____ ____ ____ ____ ____ ____ ____

frOO

_________ p__

) ________ ________ ________ _________ _________ _________ _________ ________ ________ _________ _________ ________8,

Engine speed, N, rpm

(d) Relation between air flow and engine speed.

Figure 7.- Continued. Effect of tail-pipe-nozzle area on per-formance characteristics of 195-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; sImulated altitude, 10,000 feet.

.000 12000 14.000 1000 l.000

NACA PM No. E8J28b

45

5.5

4) U)

'3 -5.0 .0 '-4

;4.5

4.0 U,

4)

43

a'3.5

0

V a) (A

•0

0 .4 43 0.

2.5

1.5 ____ ____ ____ ____ ____ ____ ____ ____ ____ ____ ____ 8,000 10,000 12,000 14,000 16, 00 18,000

Engine speed, N, rpm

Ce) Relation between specific fuel consuption based on net thrust and engine speed.

Figure 7.- Continued. Effect of tail-pipe —nozzle area on per-formance characteristics of 193-8 turbojet engine. Free-stream ram-pressure ratio, 1.11; simulated altitude, 10,000 feet.

____ ____ ____ ____il-pip

nozzle area (sq in.)

0 135

0 106

____ ____\

____ ____ ____

I \\\

46 NACA RM No. E8J2Bb

Ta 1. 1-pipe - _____ _____ _____ _____ _____ nozzle _____ _____ _____

area (sq in.)

0.135

I 0 106

S.'

43

'-4

0.

7.0

6.0

43

C

C 0

0 4.0

In

C 0

42 0. 3.0

1.0 8,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm

Ct) Relation between specific fuel consumption based on net thrust horsepower and engine speed.

Figure 7.-. Continued. Effect of tail-pipe—nozzle area on per-formance characteristics of 193-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude, 10,000 feet.

NACA PM N. E8J28b

47

5

4) U)

.0 '-4

-.

.0

4)3 w

0

V 4)

(13 .0

0 .4 4) 0. E

(0

0 C)

-4

.4 '4 '.4 C)

0. (I,

1.

Net thrust, F, lb

(g) Relation between specific fuel consumption based on net thrust and net thrust.,

Figure 7.- Continued. Effect of tail-pipe-nozzle area on per-formance characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude ,, 10,000 feet.

48

NACA RM No. EBJ28b

Tall-pipe--nozzle ____ ____ area (sq in.)

0 135

o 106

.0l

.016

.014

'4

0 -.4

.012

i ) _____ ___

1 -4

/-.01C

-/

I-

• OOE - - ______

• 006 8,000 10,000 12,000 14,000 16,000 18,000

Engine speed, N, rpm

(h) Relation between fuel-air ratio and engine speed.

Figure 7.- Concluded. Effect of tail-pipe —nozzle area on per-formance characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude, 10,000 feet.

NACA PM No. E8J28b

49

1 40C

1 20C

100C

0

Ii. 800

Simulated _____ altitude _____ _____ _____ _____

(ft)

0 5,000 o 15,000

25,000

iii _ ___

_ iii _

----v____ ___

U)

600 a)

V a>

0 a)

400 0

200

8,000 10,000 12,000 14,000 16,000 18,000 20,000 Corrected engine speed, N/J rpm

(a) Relation between corrected jet thrust and corrected engine speed.

Figure 8.— Effect of altitude on generalized performance charac-teristics of 196-8 turbojet engine. Free—stream ram—pressure ratio, 1.013. Tail —pipe—nozzle area, 135 square inches. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

i0

4,

I

80

60

50

NACA PM No. E8J28b

12 C

20

8,vu .LuuvU i,vuu i',uu'.i i,uuu i,uvu u,uO0 Corrected engine speed, wI4, rpm

(b) flelation between corrected net thrust and corrected engine speed.

Figure 8. — Continued. Effect of. altitude on generallz.ed per f'orinanee characteristics of 19B-8, turbojet engine. Free—stream ram—presswe ratio, 1.013. Tai1 —pipe-nozzle area, 135 square inches. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

aaa t 1% antS Sfl iS PItS S S - -

____ ____ ____Simulated alti tude

(f't)

0 5,000 o 15,000 O 25,000

____

o_ _ __

_ _ 7._ _ /

__7z_-__--_

NACA RM No. E8J28b

5'

w

Simulated

____ altitude ____ ____ (ft)

0 5,000 o 15,000

25,000

'I__II2____ _ 7 _ _ _ ----7

7

000 10.000 12.000 14.000 1E000 OOO 2flC

0 _____ _____ _____ _____ _____

8, -

3c

2

2C

015

L

a)

10 a)

0 U

5

Corrected engine speed, N/ rpm (c) Relation between corrected air flow and corrected eng!ne

speed.

Figure 8.- Continued. Effect of altitude on generalized per-førmanee characterIstIcs of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.013. Tail-pipe-nozzle area, 135 square inches. Performance parameters corrected to MACA standard atmospheric conditions at sea level.

0o

52

NACA RM No. E8J28b

_ I _

)0

S1mua ted altitude

(ft.)

0 5,000 o 15,000

25,000

____

___ _/

:i

20C

1 SC

1 6(

. 14

12

C 0

p. E

c 10 0 (.1

-1

0)

0 0)

0 L)

6

4 8,000 10,000 12,000 14.000 16,000 18,000

Corrected engine speed, M/f, rpm (d) Relation between corrected fuel consumption and corrected

engine speed.

Figure 8. — Continued. Effect of altitude on generalized per-formance eharacterlsttcs of 19B-8 turbojet engine. Free—stream ram—pressure ratIo, 1.013. Tail —pipe—nozzle area, 135 square inches. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

20,000

NACA RM No. E8J28b

4-) U)

4)

.0 rl

53

Simulated ____

alti tude ____ ____ (ft)

o 5,000 ___-o 15,000

25,000

_ iii__

--\__ ___ ___\N_ ___ iii _ _ _

u-OO 10.000 12.000 1400( 1n 1 flCfl or t•

.0 _-I

'4 4.._.

4) 0)

4.0 4) 4)

0

•0 4) U) • 4)

C 0 -I 4-) 0.

F.4 4,

. C.)

-4

-I C-) 4) 0. a)

. 2.0 4,

0 4)

0 ° 1.5 8 .0- - - --.--

Corrected engine speed, N/, rpm (e) Relation between corrected specific fuel consumption based

on net thrust and corrected engine speed. Figure 8.- Continued. Effect of altitude on generalized per-formance characteristics of 19B-8 turbojet engine. Free-stream rain-pressure ratio, 1.013. Tail-pipe-nozzle area, 135 square inches. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

NACA RM No. E8J28b

Simulated - altitude _____

(ft)

0 5,000 ____ - 0 15,000

o 25,000

4.0

I 3.5

3.0

2,5

20

1.50 200 400 600 800 1000 1OO

Corrected net thrust, F/6, lb Ct) Relation between corrected speciflc fuel consumption based on

iiet thrust and corrected net thrust, Figure 8.- Continued. Effect of altitude on generalized performance

characteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.013. Tail-pipe-nozzle area, 135 square inches. Perform-ance parameters corrected to NACA standard atmospheric conditions at sea level.

54

4)

1-

4)

'-4

L

'-I. 5.0

'4 4.5

4)

I,

• .018

a 0 .,.4

4)

'-4

r-4

41)

'4

V 41) 4)

0

0

.016

.014

.012

NACA RM No. E8J28b

55

.010 8,L _, _,

Corrected engine speed, rpm (g) Relation between corrected fuel-air ratio and corrected

engine speed.

Figure 8,- Concluded. Effect of altitude on generalized per-formance characteristics of l9B-8 turbojet engine. Free-stream ram-pressure ratio, 1,013. Tail-pipe-nozzle area, 135 square inches. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

bW.I.r. !IiTs 1WiaT•

H _ ':4 _r:i_

_____UI

I

56

NACA RM No. E8J28b

- Corrected engine speed, N/ft rpm (a) Relation between corrected jet thrust and corrected engine

speed. Figure 9,- Effect of altitude on generalized performance charac-

teristics of 19XB-1 turbojet engine. Static test conditions. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

NACA RM No. E8J28b

57

30

25

C.)

a) U)

r4

20

Cd

15 0

'-I

'4

•r4

('S

, 10 4)

C)

0 C.)

5

0

H Simula ted alt I tu d e

(ft) o 5,000 o 10,000

15,000 L 25,000 V 30,000

____ ____

____ ____ ____ ____ ____

____

____ ____ ____

0

___ V ______

A _________

7,000 9,000 11,000 13,000 15,000 17,000 Corrected engine speed, N/fW rpm

(b) Relation between corrected air flow and corrected engine speed.

Figure 9. — Continued. Effect of altitude on generalized per-formance characteristics of 19XB-1 turbojet engine. Static test conditions. Performance parameters corrected to NACA standard atmospheric conditions at sea level•

20CC

180C

160C

.0 r4

'4

0

140C

120C

1000

800

58

NACA RM No. E8J28b

400--'--- -.--- --.--

Corrected engine speed, N/W, rpm Cc) Relation between corrected fuel consumption and corrected

engine speed.

Figure 9.- Continued. Effect of altitude on generalized per-formance characteristics of 19XB-1 turbojet engine. Static test conditions. Performance parameters corrected to NACA standard atmospheric coziditions at sea level.

•i•i.

_

I__ It__ ____

_____________Z1 nj -

--in _

P ___

-U

NACA RM No. E8J28b 59

3.

3.

2.

2.'

1.

7,000 9,000 11,000 13,000 15,000 17,000 Corrected engine speed, N/ rpm

(d) Relation between corrected specific fuel consumption based on jet thrust and corrected engine speed.

Figure 9.— Continued. Effect of altitude on generalized per-formance characteristics of 19X8 —1 turbojet engine. •Static test conditions. Performance parameters corrected to NACA standard anospheric conditions at sea level.

.4., 4)

43

.0

a 43 U)

43

4, w

C 0 V a) (6 0

0

__ __ __ __ __ __ __L__ __

r

__Simulated altitude

( ft )

0 5,000 0 10,000 G 15,000

25,000 V 30,000

____ ____ ____

____ ____

____ ____ ____

____ ____

5\\I

-\_________ o

__ N _ ____ 5__ _ _ _ _

0 _____ _____ _____ _____ _____ _____ _____ _____ _____ _____ _____

60

NACA RM No. E8J28b

'a I .0

I 43

•0 ., 4) .).

0 •1-4

2.5 U)

0 C)

rl

0

'I -4 0 4)

U)

V 4)

4) C) 4)

0 o

1

Simulated altitude

___ (ft) ___ ___

0 5,000 o 10,000 G 15,000 --

25,000 V 30,000

iiiI_ _ _ iiii__i__iii - - L= --_ _

00 300 500 700 900 1100 1300 Corrected jet thrust, F/5, lb

Ce) Relation between corrected specific fuel consumption based on jet thrust and corrected jet thrust,

Figure 9.— Continued.. Effect of altitude on generalized performance characteristics of 19XB—]. turbojet. engine. Static test conditions. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

NACA RM N. E8J28b

61

.02C

• 010 ____ ____ ____ ____ ____ ____ ____ ____ ____ ____ ____

7,___ --.-- -, Corrected engine speed, N/ rpm

(f) Relation between corrected fuel-air ratio and corrected engine speed.

Figure 9.- Concluded. Effect of altitude on generalized per-formance characteristics of 19XB-1 turbojet engine. Static test conditions. Performance parameters corrected to NACA standard atmospheric conditions at sea level0

DOO 9...00o 1rnn isrw i'i t--r

4-4

0

.016

a .014 '4

a, 4) C) a,

.012 0

C.)

Simula t.e d altitude

(ft)

o 5,000 o 10,000 G 15,000

25,000 V 30,000

___ ____-

____ ____ ____

___ ___ ___ ___

____ ____ ____

___ ___

>

7.

N________ / - -0

62

NACA RM No. E8J28b

18CC

160C

1 40C

.D 120C r1

100C U)

4)

8CC

0) 4)

0 a)

060C

40C

200 ____ ____ ____ ____ ____ ____

8, LVU.'U .L,UOU 14,000 16,000 18,000 Corrected engine speed, N/I, rpm

(a) Relation between corrected jet thrust and corrected engine speed.

Figure 10.- Effect of ram-pressul'e ratio on generalized per-formance characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

-- - -

Ow

Free-s tream ram-

• pressure ratio

0 1.012 o 1.122 o 1.155

____ ____ ____ ____ ____ ____

____/ ___ ___

-

Free-stream ram-

pressure ratio

0 1.012 0 1.122

1.155

__ __ __

/1 --ciiii___ -°

_ __

____7__

,000 10,000 12.000 14.000 16.000 1flCfl

14(

12(

1O(

0 '-4

BC

4.)

I 6C

20

E

NACA RM No. E8J28b 63

Corrected engine speed, N/-1, rpm (b) Relation between corrected net thrust and corrected engine

speed.

F!gure 10,- Continued. Etfect of ram-pressure ratio on gener-alized performance characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

NACA RM No. E8J28b 64

D

35

30

Free-stream ram-

pressure ratio

0 1.012 o 1.122 o 1.155

25 0

5

01 I I I I I I I I I I 8,000 10,000 12,000 14,000 16,000 18,000

Corrected engine speed, N/'f, rn (e) Relation between corrected air flow and corrected engine

speed. Figure 10.- Continued. Effect of ram-pressure ratio on gener-alized performance characteristics of 19B-8 turbojet engine. Tail-pipe--nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters eorrected to NACA standard atmospheric conditions at sea level.

a

a 0

I

NACA RM N. E8J28b

65

,uoo 10,000 12,000 14,000 16,000 18,000 Corrected !nglne seed, N/-(, rpm

(d) Relation between corrected fuel consumption and corrected engine speed.

Figure 10.- Continued. Efrect of ram-pressure ratio on gener-alized performance characteristics of l9B-8 turbojet engine. Tail-pipe--nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to MACA standard atmospheric condition! at sea level.

5—

____-..

____ ____

Free—s tream ram -

pressure ratio

0 1.012 0 1.122 c 1.155

:'_

-4 ____

-\-_ ____—_c

_

•__ 1 __

4.

4.1

2 .

2.(

1c

66

NACA RM No. E8J28b

4) U)

'I

•0 '-I

4

4) U)

54

4)

4)

0

V U,

0 4) a

8,000 10,000 12,000 14,000 16000 18,000 0Corrected engine speed, N/Ië, rpm

(e) Relation between corrected s pecific fuel consumption based on net thrust and corrected engine speed.

Figure 10. — Continued. Effect of ram—pressure ratio on gener-alized performance characteristics of 195-B turbojet engine. Tail—pipe--iozz1e area. 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

Ii _________________ HL Free—stream

ram—pressure ratio

0 1.012 o 1.122 O 1,155

_________

C

4.

4.0

3.5

3.0

2.5

2.0

1.5

1.0

NACA RM No. E8J28b

67

-

U)

0 I-I

-%

'C

'a

43 U)

43

4)

0

'Ti

II)

.0

0 -4 4.) 0.

0 W0 bOO 800 1000 1200 1400 Corrected net thrust, PJ6, lb

(f) Relation between corrected specific fuel consumption based on net thrust and corrected net thrust.

Figure 10. — Continued. Effect ofrarn—pressure ratio on generalized performance characteristics of 198-8 turbojet engine. Tail—pipe-nozzle area, 106 square inches; simulated altitude; is,000 feet, Performance parameters corrected to NACA standard atmospheric conditions at sea level1

68

NACA RM No. E8J28b

4)

•0

8.0 a.

i 7.0 0)

0 0. 0) U)

0

6.0 4) I

5.0

0

•0 0) U)

o 4.0 0

-.4 4) 0.

3.0

Free-stream ram -

p re s sure ratio

o 1.012 o 1.22 ' 1.155

4)

C) -4 '4 -.4 0 0, 0.

U,

1.0 I I I I I I I I I

8 1AflAU 1D (Yfl 1A A(V 1 nAn 1Q AAA

Corrected engine speed,. N/;rpm (g) Relation between specific fuel consumption based on net

thrust horsepower and corrected engine speed. Figure 10,- Continued. Effect of ram-pressure ratio on gener-

alized performance characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

NACA RM No. •E8J28b

69

.022

.02C

.016

'4

:1 .016

.014 '4

4)

4)

.012 0 C)

.010

.008 ____ ____ ____ ____ 8,000 10,000 12,000 14,000 16,000 18,000

Corrected engine speed, N/-a, rpm (h) Relation betweencorrected fuel-air ratio and corrected

engine speed. Figure 10.- Concluded. Effect of ram-pressure ratio on gener-

alized performance characteristics 3f 19B-8 turbojet engine. Tail-pipe-nozzle area, 106 square inches; simulated altitude, 15,000 feet. Performance parameters corrected to NACA standard atmospheric conditions at sea level.

'-11 Free-stream

ram-pressure ratio

0 1.012 a 1.122

1.155

____ __L ____

60

-

40

20

____ ____ ____ ____ ____ ____ ____ ____SImulated alt Itude

(ft) 0 5,000 0 20,000

30,000

____

I / - - 1

0 __ __ __ __ __ __ __ __ __ __ __

ii

I

1:

,-1

C

a

(6

.Tj

'-4 '-I '-4 E V

--4

V 4) C,

0 C.)

70

NACA PM No. E8J28b

0 100 200 300' 400 500 Corrected true airspeed, V'0/1,' mph

(a) Relation between corrected windmilling drag and corrected true airspeed. Windinilling drag and true airspeed corrected to NACA standard atmospheric conditions at sea level,

Figure 11,- Effect or altitude on w'indmilllng drag character-istics of 19B-8 turbojet engine, Tail-pipe-nozzle area, 135 square inches,

_____ _____ _____Simulated alt i tude

(ft)

0 5,000

0 20,000

0 30,000

-- /

-.

.15

4, tou) (

4,

H .05

NACA PM to. E8J28b

7

C, 4)

0

(0

0

'4

(1

V 4)

43 C, 4)

0 C)

0 100 200 300 400 500 True airspeed, V, mph

(b) Relation between true airspeed and ratio of windmilling drag to maximum net thrust.

A __ A' _

__IEI___

0 1000 2000 3000 4000 5000 601 Corrected engine windmilling speed, N/i, rpm

(c) Relation between corrected air flow and corrected engine wind-milling speed. Air flow and engine windmilling speed corrected to NACA standard atmospheric conditions at sea level.

Figure 11.- Concluded. Effect of altitude on wthdmilltng drag characteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches.

15

10

5