restriction/classification cancelled/67531/metadc...19b-21 19b-8, and 19-xb-1 jet-propulsion engines...
TRANSCRIPT
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Restriction/Classification Cancelled
Restriction/Classification Cancelled
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6w,
NACA RM No. E7C 13
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
RESEARCH MEMORANDUM
for the
Bureau of Aeronautics, Navy Department
ALTITUDE-WIND-TUNNEL INVESTIGATION OF THE
19B-2 1 19B-8, AND 19-XB-1 JET-PROPULSION ENGINES
III - PERFORMANCE AND WINDIMILLING DRAG CHARACTERISTICS
By William A. Fleming and Robert 0. Dietz, Jr.
SUIRAARY
The performance characteristics of the 19B-8 and 19XB-1 turbo-jet engines and the windnilling-drag characteristics of the 19B-6engine were determined in the Cleveland altitude wind tunnel. Theinvestigations were conducted on the 19B-8 engine at simulatedaltitudes from 5000 to 25 .,000 feet with various free-stream ram-pressure ratios and on the 19XB--1 engine at simulated altitudesfrom 5000 to 30,000 feet with. approximately static free-streamconditions.
Data for these two engines are presented to show the effect ofaltitude, free-stream ran-pressure ratio, and tail-pipe-nozzle areaon engine performance. A 21-percent reduction in tail-pipe-nozzlearea of the 19B-8 engine increased the let thrust 43 percent th.enet thrust 72 percent, and the fuel consumption 64 percent. Anincrease in free-stream ram-pressure ratio raised the jet thrust andth.e air flow and lowered the net thrust through.out the entire rangeof engine speeds for the 19B-8 engine. At similar operating condi-tions o, the corrected jet thrust and corrected air flow were approxi-mately the same for both engines, and the corrected specific fuelconsumption based on 'let thrust was lower for the 19XB-1 engine thanfor the 19B-8 engine. The thrust and air-flow data obteined withboth. engines at various altitudes for a. given free-stream ram-pressure ratio were generalized to standard sea-level atmosphericconditions. The performance parameters involving fuel consumptiongeneralized only at high. engine speeds at simulated altitudes ashigh as 15 ,,000 feet. The windmilling drag of th.e 19B-8 engineincreased rapidly as the airspeed was increased.
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INTRODUCTION
Investigations have been conducted in the Cleveland altitudewind tunnel at the request of the Bureau of Aeronautics, NavyDepartment, to determine the performance and operational character-istics of the 19B-2, 19B-8, and 19XB•-1 turbojet engines. Opera-tional characteristics of all three engines are presented in refer-ence 1 and the performance characteristics of the turbine in the19B-8 engine are presented in reference 2. Performance character-istics of the 19B-8 and 19XB-1 turbojet engines over a range ofsimulated flight conditions are presented herein. Data for the19B-2 engine are omitted because it is an earlier production modelof the 19B-8 engines
The effects of altitude, free=stream ram-pressure ratio s andtail-pipe-nozzle area on engine performance are shown, The perform-ance data are generalized to show the applicability of methods usedto determine performance at any altitude from rtiuns at a givenaltitude. Windmilling drag characteristics of the 19B-8 e:Lgine arealso Dresentedo
Results are shown for the 19B-8 engine at simulated altitudesfrom 5u00 to 25,000 feet at Free-stream ram-pressure ratios uy to1.155 ana for the 19XB-1 engine at simulated altitudes from 5000 to30 .,000 feet at free-stream ram-pressure ratios below 1,01.
DESCRIPTION OF :ENGINES
19B- 8 engine. - The 19B-8 turbojet engir :. has a static sea-level rating of 1365 pounds of thrust at an engine speed of17,500 rpm,, At this rating the engine has an air flow of 28 poundsper second and a fuel consumption of approximately 1800 pounds per
hour. The over-all length. of the engine is 1042 inches with. the
adjustable tail cone extended; the maximum diameter is 204 nches
and the total weight is 825 pounds, The maximum diameter does notinclude the accessory group mounted on the front bearing sir portat the compressor inlet.
This engine has a six-stage axial-flow compressor, an annular-type combustion chamber, and a single-stage turbine. Fuel isinjected downstream into the combustion chamber through 24 nozzlesmounted circumferentially in the manifold at the forward end of thecombustion chamber. The gases leaving the turbine pass through a
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tail pipe equipped with a movable inner cone, which has a totalaxial travel of 5 inches. When the inner cone is moved from the"t in, 11 or forward osition, the ta
i
., position to the "4-inches-out" p l-
- Lpipe-nozzle area is decreased from 135 to 106 square inches. Thechange in area is nearly linear with respect to the travel of thetail cone. In order to cool the oil for lubricating the enginebearings and accessory drive, a cylindrical oil cooler 2 feet longwith an inside diameter of 14 inches is attached to the front ofthe engine. The inner wall of the cylinder provides the lieat-exchanger surface for the oil.
19XB-1 engine. - The 19XB-1 turbojet engine is a modificationof the 19B-8 engine that contains a ten-stage axial-flow compressorand has the same over-all engine dimensions. The tail - pipe lias afixed inner cone and the tail—,pipe-nozzle area is 103 square inches.The 19XB-1 engine has a sea-level rating of approxim pately 1400 poundsstatic thrust at a rated engine speed of 16,500 rpm. At this ratingthe air flow is 29 pounds per second and the fuel consumption isapproximately 1600 pounds per hour.
INSTALLATION AND PROCEDURE
Each engine was installed in a nacelle mounted under a stubwine that extended from one side of the 20-foot-diameter test sec-tion of the wind tunnel. (See fig. 1.) The engine air was admittedat the front of the nacelle, and a small duct beneath. the leadingedge of the wing admitted air to cool the outer well of the combus-tion chamber and the tail pipe. For the runs at approximately staticconditions, when the free-stream ran-pressure ratio was *b e low 1.01,the part of the cowling enclosing the tail pipe and the rear of thecombustion chamber was removed because the cooling-air flow throughthe cowling was inadequate.
Survey rakes were installed at several stations in the engine(fig, 2) in order to obtain temperature and pressure measurementsfrom which the component and over-all performance of the enginecould be determined. The methods used to calculate performanceparameters are presented in the appendix.
Throughout the part of the investigation discussed, a 6-mesh.screen of 0.032-inch-diameter steel wire was attached to the frontflange of the oil cooler to protect the compressor from flyingobjects in the tunnel.
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The 19B-8 engine was operated at simulated altitudes from 5000to 25,000 feet with. free-stream ram-pressure ratios up to 1.155; andthe 19KB-1 engine, from 5000 to 30,000 feet at approximately staticfree-stream conditions. The tunnel temperature was held at approxi-mately NACA standard values for each flight condition. At eachsimulated flight condition, the engine was run over the full rangeof operable engine speeds, The fuel specified for these engines,62-octane unleaded gasoline, was used.
During the investigation, the combustion chambers were modifiedseveral times (see reference 1) in an attempt to increase theoperating range of the engines at high altitudes. Because none ofthe modifications made any significant improvement in the operatingrange of the engine, all the data presented are for the engines withthe standard combustion chambers as received from the manufacturer.
SYMBOLS
The following symbols are used in this paper:
A cross-sectional area, square feet
B t^^rust sca e reacting, pounds
CDinstallation drag coefficient
C specific heat of gas Et constant pressure, Btu perpound oR
D windmilling drag, pounds
Fj jet thrust, pounds
Fnnet thrust, pounds
f /a fuel-air ratio
g acceleration of gravity, 32.2 feet per second per second
i mechanical equivalent of heat, 778 foot-pounds per Btu
m mass flow, slugs per second
N engine speed, rpm
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P total pressure, pounds per square foot absolute
p -static press-Lire, pounds per square foot absolute
R gas constant, 53.4 foot-pounds per pound OR
S wing area, square feet
T total temperature ., OR
Ti indicated temperature, OR
t static temperature, OR
thp net thrust horsepower
V velocity, feel per second
VT velocity, miles per hour.
Waair flow, pounds per second
Wf fuel consumption, pounds per hour
WflFj specific fuel consumption based on jet thrust, poundsper hour pound thrust
Wf/Fn specific fuel consumption based on net thrust, poundsper hour pound thrust
Wf/thp specific fuel consumption based on net thrust horse-power, pounds per hour per thrust horsepower
Y ratio of specific beats for gases
5 ratio of absolute tunnel static pressure to absolutestatic pressure of NACA standard atmospheric con-ditions at sea level, -p0/2116
ratio of absolute tunnel static temperature toabsolute static temperature of NACA standardatmospheri c conditions at sea level, t0/519
P mass density, slugs per cubic foot
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Subscripts:
a air
f fuel
9 gas
station at which static pressure in jet reaches ambientstatic pressure
0 ambient free stream
1 cowl inlet
2 compressor inlet
5 turbine outlet
The follo ,,,AnG, performance parameters generalized to NACA standardatmospheric conditions at sea level were used
DW/F) corrected.. t17indmm.11ing dra ,, pounds
FO/5 corrected jet thrust, pounds
rV6 corrected net thrust, pounds
(f/a)/e corrected fuel-air ratio
I VO_N/ corrected -engine speed J rpm
V' 0/ corrected true airspeed, miles per hour
Wa' F(913 corrected-air flow, pounds per second
corrected fuel consumption, pounds per hour
Wf / (F P e) corrected specific fuel consumption based on jet thrust,pounds per hour per pound thrust
Wfl(FnVO—) corrected specific fuel consumption based on net thrust,,Dounds per hour per pound thrust
Wf/tbp corrected specific fuel consumption based mi net thrusthorsepower, pounds per hour per thrust horsepower
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PERFORMANCE CHp.RACTERISTICS
Effect of Inlet Screen
The screen installed at the cowl inlet for the wind-tunnelinvestigation caused a total-pressure loss between the free streamand the compressor inlet that varied with the corrected enginespeed. The ram-pressure ratios used in this report are based onmeasurements of free-stream total pressure instead of compressor-inlet total pressure. Because the-tunnel-test-section airspeed wasused to calculate net thrust and net thrust horsepower, the engine-performance characteristics involving these parameters were slightlyaffected. Th.e.relation between corrected engine speed and the ratioof compressor-inlet total pressure to free-stream total pressure isshown in figure 3. The curve shown in this figure was determinedfrom data obtained throughout the entire range of flight conditionsinvestigated.. The compressor-inlet ram-pressure ratio at any enginespeed may be obtained for the data presented by multiplyi thefree-stream ram-pressure ratio by the value of P2/PO obtained fromfigure 3 for the corresponding corrected engine speed.
Effect of Altitude
19B-8 engine, - The 19B-8 engine performance at simulatedaltitudes of 5000, 15,000, and 25 9 000 feet, a free-stream ram-pressure ratio of 1,013, and with a tail-pipe-nozzle area of135 square inches, is shown in figure 4. Because of the high. min-imum speed of the engine at altitudes above 17,000 feet (refer-ence 1), no data could be sho —vn for an engine speed, below15,500 rpm at an altitude of 25.000 feet, The variations in jetthrust, net thrust, fuel consumption, air flow, and specific fuelconsumption based on net thrust with increasing altitude are s;olmin figures 4(a) to 4(e), respectively,
Above an engine speed of 13,000 rpm the fuel-air ratios atsimulated altitudes of 5000 and 15,000 feet were the same (fig. 4(f)).At an altitude of 25,000 feet, however, the fuel-air ratio through.- .out the entire operable range of the engine was higher tLa.n at5000 and 15,000 feet. The high. fuel-air ratio obtained at25,000 feet was probably related to the reduction in combustionefficiency. A low combustion efficiency at a simulated altitude of25,000 feet was indicated by the narrow operating range of theengine as limited by combustion blow-out. At engine speeds below
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13,000 rpm, the increase in fuel-air ratio at an altitude of15,000 feet above that at 5000 feet was also caused by the reductionin combustion efficiency that accompanied an increase in altitude.
19XB-1 engine. - The 19XB-1 engine performance at approximatelystatic conditions (free-stream ram-pressure ratio less than 1.01)and simulated altitudes of 5000 to 30,000 feet is shown in figure 5.Maximum engine speed was Limited by high. tail-pipe temperatures andabove an altitude of 15,000 feet the minimam engine speed began toincrease. The reduction in jet thrust, fuel consumption, and airflow that resulted from an. increase in altitude, is shown infigures 5(a) to 5(c), respectively. No net-thrust curve is presentedinasmuch as these data were obtained at approximately static condi-tions where the jet and .net thrusts are equal. Variation of specificfuel consumption based on jet thrust with altitude and engine speedis shown in figure 5(d). For high. engine speeds, the s pecific fuelconsumption was slightly higher at altitudes of 25,000 and30,000 feet than at the lower altitudes. The effect of altitude onthe relation between engine speed and fuel-air ratio is shown infigure 5(e).
Effect of Ram-Pressure Ratio
The effect of varying the free-stream ram-pressure ratio from1,012 to 1.155 throug,out the operable speed range of the 19B-8engine with a tail-pipe-nozzle area of 106 square inches at analtitude of 15,000 feet is presented in figure 6. The maximumengine speed obtainable with. the tail cone 4 inches out (tail-pipe-nozzle area, 106 sq in.) was limited to about 16,500 rpm by hightail-pipe temperatures. Increasing the ram-pressure ratio raisedthe jet thrust (fig. 6(a)) and the air flow (fig. 6(c)) and loweredthe net thrust (fig. 6(b)) throughout the entire range of enginespeeds. At all engine speeds below 16,500 rpm the fuel consumptionwas lowered as the ram-pressure ratio was increased (fig. 6(d)).
At low engine speeds, the specific fuel consumption based onnet thrust increased as the ram-pressure ratio was raised from 1,012to 1.155 (fig, 6(e)). Above a speed of 14,000 rpm, however, thespecific fuel consumption increased only for a change in ram-pressure ratio from 1.012 to 1.122, The specific fuel consumptionbased on net thrust horsepower decreased with. increasing ram-ypressure ratio throughout the entire range of engine speeds(fig. 6(f)). At an engine speed of 16,500 rpm, the specific fuel
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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(fig. 6(e)) and the specific fuel consumption based on net thrusthorsepower'decreaeed from 5.25 to 1.85 for a corresponding increasein rasa-pressure ratio (fig. 6(f)). The relation between specificfuel consumption based on net thrust and net thrust is shown infigure 6(g).
The fuel-air ratio (fig, 6(h)) was reduced throughout theentire range of engine operation as the ram-pressure ratio wasincreased.
Effect of Tail-Pipe-Nozzle Area on Performance
The effect of a 21-percent reduction in tail-pipe-nozzle area(from 135 to 106 sq in.), which was obtained by moving the tailcone from the in position to the 4-inches-out position, on theperformance of the 19B-8 engine at a free-stream ram-pressure ratioof 1.131 and at a simulated altitude of 10,000 feet, is shown infigure 7. With a tail-pipe-nozzle area of 106 square inches, themaximum engine speed was limited to 16,500 rpm by the turbine-inlettemperature (1500 0 F). The reduction in tail-pipe-nozzle areaincreased the jet thrust (fig. 7(a)), net thrust (fig. 7(b)), andfuel consumption (fig. 7(c)) and lowered the air flow (fig. 7(d))throughout the range of engine speeds. At an engine speed of16,500 rpm, the reduction in nozzle area increased the ,het thrust43 percent (fig. 7(a)), the net thrust 72 percent (fig. 7(b)), andthe 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 fuelconsumption based on net thrust horsepower (fig. 7(f)) were lowerwith a tail-pipe-nozzle area of 106 square inches than with anarea of 135 square inches. At an engine speed of 16,500 rpm, thespecific fuel consumption based on net thrust was approximately6 percent higher with a tail-pipe-nozzle area of 135 square inchesthan with an area of 106 square inches. Changing the tail-pipe-nozzle area had no effect on the relation between the specific fuelconsumption based on net thrust and the net thrust (fig. 7(g)).
The reduction in tail-pipe-nozzle area increased the backpressure on the turbine and as a result the engine required a higherturbine-inlet temperature, that is, a higher fuel-air ratio to main-tain a given engine speed. The increase in fuel-air ratio resultingfrom a reduction in tail-pipe-nozzle area is shown in figure 7(h).
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GENERALIZED PERFORMANCE
The concept of flow similarity and the application of dimen-sional analysis to the performance of turbojet engines has led tothe development of the pressure and temperature reduction factors Sand e with which data obtained at several altitudes may be gener-alized. Data thus far presented that were obtained at various alti-tudes and free-stream ram-pressure ratios are generalized to NACAstandard conditions at sea level in order to test the applicabilityof the factors 8 and 8 to the performance of the 19B-8 and 19XB-1turbojet engines.
Application of the factors 8 and e give the following gen-eralized performance variables; corrected engine speed, N,1;corrected jet thrust, F j18; corrected net thrust, Fn/b; cor-rected air flow, (W,10)l5, corrected fuel consumption, Wf/(&/O);corrected specific fuel consumption based on jet thrust, Wf1(FjvfG_);corrected specific fuel consumption based on net thrust, Wf/(Fn,/O-);
corrected specific fuel consumption based on net thrust horsepower,Wf/thp; and corrected fuel-air ratio, (f/a)/8.
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 theworking fluid included in the dimensional analysis. This exclusionadmittedly lessens the possibilities of successfully generalizingthe data in cases where the combustion efficiency changes.
The effect of altitude on the generalized performance of the19B-8 and 19XB-1 engines is shown in figures 8 and 9, respectively.The corrected jet t"_=st (figs, 8(a) and 9(a)), the corrected netthrust (fig. 8(b)), and the corrected air flow (figs. 8(c) and 9(b))generalized to a single curve for each engine. The parametersinvolving fuel consumption, that is, corrected fuel consumption(figs. 8(d) and 9(c)), corrected specific fuel consumption(figs. 8(e), 9(d), and 9(e)), and corrected fuel-air ratio(figs. 8(g) and 9(f)), generalized only in the upper range of enginespeeds for simulated altitudes as high as 15,000 feet.
Performance data obtained at various ram-pressure ratios didnot generalize to a single curve inasmuch as the factors 8 and 0are not meant to correct for the changes in cycle efficiency thataccompany variations in raga-pressure ratios. The effect of ram-pressure ratio on the generalized performance of the 19B-8 engine
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is shown in figure 10. The performance parameters involving cor-rected 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 netthrust (figs. 10(e) and 10(f)), specific fuel consumption based onnet thrust horsepower (fig. 10(g)), and corrected fuel-air ratio(fig. 10(h) ) produced a family of curves for various free-streamram-pressure ratios.
In the use of generalized performance curves, the effect ofram-pressure ratio on engine performance at various altitudes canbe determined only by use of the entire family of curves. It isimpossible to predict the performance parameters involving fuelconsumption at low engine speeds or at high altitudes because theeffects of varying combustion efficiency are excluded from thefactors used in this report.
A comparison of generalized performance data, of the 19B-8 and19XB-1 engines at a corrected engine speed of 17,000 rpm is pre-sented in the following table:
Engine
19B- 8 . 19XB-1Tail-pipe-nozzle area, sq in. 106. 103Corrected engine speed, NI'le-, rpm 17,000 17,000Corrected jet thrust, F l /b, lb 1370 1375Corrected air flow, Wa T15, lb/sec 27.0 26.5Corrected fuel consumption, Wf /NFe, lb/hr 1965 1825Corrected specific fuel consumption based on 1.435 1.328
jet thrust, Wf l(F j ^,105), lb/(hr)(lb thrust)Corrected fuel-air ratio, Wa Ve 0.0202 0.0192
Data presented in this table were obtained from figure 9 forthe 19X-B-1 engine at static conditions and from figure 10 for the19B-8 engine at a free-stream ram-pressure ratio of 1.012. Atstatic conditions, the ejector effect of the jet produced a velocityin the tunnel test section at a maximimi engine speed equivalent toa free-stream ram-?pressure ratio of approximately 1.012. For theconditions given in the table, the compressor-inlet ram-pressureratio for both engines was 0.938. Generalized data are used becausethe inlet-air temperatures were different for the two engines.
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The corrected jet thrust and the corrected air flow of bothengines at a corrected engine speed of 17,000 rpm were approximatelythe same, The corrected specific fuel consumption based on jetthrust was lower for the 19XB-1 engine than for the 19B-8 engine,
WINDMILLING DRAG
The magnitude of the windmilling drag obtained when an engineis inoperative and is allowed to windmill during flight is equal tothe change of momentum of the air as it passes through the engine.A limited amount of windmilling drag data was obtained for the 19B-8engine at simulated altitudes from 5000 to 30,000 feet and tunnel-test-section airsLDeeds from 176 to 373 miles per hour with thescreen installed at the cowl inlet. 'Generalized values of wind-milling drag (fig. 11(a)) show that the windmilling drag increasedrapidly as the corrected true airspeed was increased. A singlecurve was drawn through the generalized values obtained at variousaltitudes. The relation between the ratio of windmilling drag tothe maximum net thrust obtainable with a tail-pipe-nozzle area of135 square inches and the tunnel-test-section true airspeed is shownin 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 was2.5 percent of the maximum net thrust obtainable at this airspeedand that at an airspeed of 375 miles per hour the windmilling dragwas 11 percent of the maximum net thrust obtainable at this air-speed, The generalized values of air flow obtained with the enginewindmilling (fig. 11(c)) fell on a single curve for all altitudesinvestigated.
SU 44ARY OF RESULTS
From an investigation in the Cleveland altitude wind tunnel ofthe performance characteristics of the 19B-8 and 19XB-1 turbojetengines and the windmilling drag characteristics of the 19B-8engine, the following results were obtained:
1. A reduction_ in tail-pipe-nozzle area of the 19B-8 enginefrom 135 to 106 square inches at a simulated altitude of 10,000 feetand a free-stream ram-pressure ratio of 1.131 gave an increase injet thrust, net thrust, and fuel consumption with a reduction in airflow and specific fuel consumption based on net thrust throughoutthe entire range of engine speeds. This reduction in tail-pipe-nozzle area at an engine speed of 16,500 rpm raised the jet thrust43 percent, the net thrust 72 percent, and the fuel consumption64 percent.
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2. With. a tail-pipe-nozzle area of 106 square inches on the19B-8 engine, an increase in free-strewn ram-pressure ratio from1.012 to 1.155 at a simulated altitude of 15,000 feet raised the jetthrust and air flow and lowered the net thrust throughout the entirerange of engine speeds. The fuel consumption was lowered at allengine speeds below 16,500 rpm.
3, At similar operating conditions, the corrected jet thrustand the corrected air flow were approximately the same for bothengines and the corrected specific fuel consumption based on jetthrust was lower for the 19XB-1 engine than for the 19B-8 engine.
4. With. th.e use of generalizing pressure and temperaturefactors, jet-thrust, net-thrust, and air-flow data obtained fromboth the 19B-8 and 19XB-1 engines at a given ram-pressure ratio andtail-pipe-nozzle area, and at any altitude could be used to estimatethe performance of the engines at any other altitude. The perform-ance parameters involving fuel consumption generalized only at highengine speeds at simulated altitudes up to 15,000 feet. Above thisaltitude these parameters did not generalize at any engine speedo
5. Th.e windmilling drag of the 19B-8 engine while it wasinoperative amounted to 2.5 percent of the maximum net thrust of theengine at an airspeed of 200 miles per hour, and 11 percent of themaximum net thrust of the engine at an airspeed of 375 miles per hour.
Aircraft Engine Research Laboratory,National Advisory Committee for Aeronautics,
Cleveland, Ohio.
6/-M^7r2y^/
William A. Fleming,Aeronautical Engineer.
Robert 0. Dietz, Jr.,Mechanical Engineer.
Approved:Fred W oun
Mech.anica Ena " r.
Abe Silverstein,Aeronautical Engineer.
rl
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APPENDIX -- M HODS OF CALCULATION
Temperature. - A sample thermocouple of the type used was cali-brated cold up to a Mach number of 0.8. This calibration showed thethermocouple measured the static temperature plus approximately85 percent of the adiabatic temperature rise due to the impact ofthe air on the thermocouple. Static temperature may be determinedfrom indicated temperature by applying this factor to the adiabaticrelation between temperature and pressure:
Ti
t1+0.85 _
p )T1
and the total temperature is
=Tit
7
T p
1+0.85 (P^7
-1
In order to detQrmine the static temperature of the jet, theassumption was made that no loss occurred in total temperature ortotal pressure between station 5 in the tail pipe and station j inthe jet. When this assumption is made, the static temperature inthe jet is
t jT5
= ^_" 5
P'5O^
The tunnel-test-section static temperature was determined fromthe indicated temperature measured in the large section of thetunnel just ahead of the test section. Because the velocity in thelarge section of the tunnel was very low, the indicated temperaturemeasured in the large section is equal to the test-section totaltemperature. The static temperature in the test section then is
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TOt0 - 70-1
PO y0
p0
Tunnel-test-section velocity. - The tunnel-test-sectionvelocity was obtained from
YO-1
_ FJ,,7 Po YO _V t0.^ p0 ^ J
Air flow, - The air flow through. the engine was determinedfrom pressure and temperature measurements obtained with foursurvey rakes at the cowl inlet (station 1, fig. 1). The equationused to calculate the air flow is
yl-1
p l Al , 2J g c_p % Pl _ ylWa = p l Al Vl g = R ti
F p, ^ - 1
Thrust and windmilling drag. - Two methods were used to cal-culate the jet thrust. For the tests at approximately staticconditions when the tunnel-test-section velocity was low, thethrust was measured directly on the balance scales. The externaldrag of the nacelle and th.e initial momentum of the inlet air wereadded to the scale thrust reading, which gave the following equa-tion for jet thrust;
F j = B + 1/2 p0 V02 CD S + ma VO(1)
Inasmuch as the scale-thrust term B of equation (1) was themajor part of the total thrust and could be measured within±2 pounds, the jet thrust could be determined accurately at lowtest-section velocities. Although. variations in the drag coeffi-cient accompanying a change in inlet-velocity ratios could not bedetermined, an error of as much. as 50 percent in the drag coeffi-cient would result in less than a 1-percent error in thrust attest-section airspeeds below 100 feet per second.
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For test-section airspeed: of 100 miles per hour or more thejet th.rust was determined by using the mass air flow measured atthe cowl inlet, the fuel consumption ., and the jet velocity at thevena contracta in the following equation:
F j = (ma + mf . ) Vj
(2)
Calculating the jet thrust with these two methods for the statictests gave a calibration factor that could be applied to equation (2)FThe value of the calibration factor, which is the ratio of measuredscale thrust to calculated tail=pipe thrust, was found to be 1,015.By expansion of equation (2) and application of this factor to it'the following expression used to calculate the jet thrust for all ofthe tests except the static tests was obtained:
7j
F j = 1.015 mg r.^ 2J g cp, j t j (T5 ) - 1
0 —^
The net thrust was then determined by subtracting the free-stream momentum of the inlet air from the jet thrust:
Fn = F j - ma VO .
The net thrust horsepower was determined from the net thrustand the true 'airspeed by the expression:
Fn VOthe = 550
The wi.ndmilling drag of the engine was determined from theproduct of the mass air flow and the change in velocity of the airthat passes through. the engine by use of the equation
Dw = ma (VO - Vj)
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REFERENCES
1. Fleming, William A.: Altitude-Wind-Tunnel Investigation ofthe Westinghouse 19B-2, 19B-8, and 19XB-1 Jet-PropulsionEngines. I - Operational Characteristics. NACA MR No.EEE06 ) Bur. Aero., 1946
2. KKK, Richard P., and Suozzi, Frank L.: Altitude-Wind-TunnelInvestigation of the.19B-2, 19B-8, and 19XB-1 Jet-PropulsionEngines. II - Analysis of Turbine Performance of the 19B-8Engine. NACA RM No. E7AO8, Bur. Aero., 1947.
I
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INDEX OF FIGURES
Figure 1. - Installation of 19B turbojet engine in underslung nacellefor investigation in altitude wind tunnel.(a) With cowling.(b) Without cowling.
Figure 2. - Longitudinal section of 19B turbojet-engine installationshowing measuring stations.
Figure 3. - Relation between total-pressure ratio across cowl-inletscreen and corrected engine speed for range of simulated altitudesand flight speeds.
Figure 4. - Effect of altitude on performance characteristics of19B-8 turbojet engine. Tail-pipe-nozzle area, 135 square inches;free-stream ram-pressure ratio, 1.013.(a) Relation between jet thrust and engine speed.(b) Relation between net thrust and engine speed,(c) Relation between fuel consumption and engine speed.(d) Relation between air flow and engine speed,(e) Relation between specific fuel consumption based on net thrustand engine speed.
M Relation between fuel-air ratio and engine speed.
Figure 5, - Effect of altitude on performance characteristics of19XB-1 turbojet engine. Approximately static conditions.(a) Relation between jet thrust and engine speed,(b) Relation between fuel consumption and engine speed.(c) Relation between air flow and engine speed.(d) Relation between specific fuel consumption based on jet thrustand engine speed.
(e) Relation between fuel-air ratio and engine speed.
Figure 6. - Effect of free-stream ram-pressure ratio on performancecharacteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area,106 square inches; simulated altitude ,, 15,000 feet,(a) Relation between jet thrust and engine speed.(b) Relation between net thrust and engine speed.(c) Relation between air flow and engine speed.(d) Relation between fuel consumption and engine speed,(e) Relation between specific fuel consumption based on net thrustand engine speed.
(f) Relation between specific fuel consumption based on net thrusthorsepower and engine speed.
(g) Relation between specific fuel consumption based on net thrustand net thrust.
(h) Relation between fuel-air ratio and engine speed.
CONFIDENTIAL
-
NACA RM No. E7C13 CONFIDENTIAL 19
Figure 7. - Effect of tail-pipe-nozzle area on performance charac-teristics of 19B-8 turbojet engine. Free-stream ram-pressureratio, 1.131; simulated altitude, 10,000 feet.(a) Relation between jet thrust and engine speed.(b) Relation between net thrust and engine speed.(c) Relation between fuel consumption and engine speed.(d) Relation between air flow and engine speed.(e) Relation between specific fuel consumption based on net thrustand engine speed.
(f) Relation between specific fuel consumption based on net thrusthorsepower and engine speed.
(g) Relation between specific fuel consumption based on net thrustand net thrust.
(h) Relation between fuel-air ratio and engine speed.-
Figure 8. - Effect of altitude on generalized performance charac-teristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio,1.013. Tail-pipe-nozzle area, 135 square inches. Performanceparameters corrected to NACA standard atmospheric conditions atsea level.(a) Relation between corrected jet thrust and corrected engine
speed.(b) Relation between corrected net thrust and corrected engine
speed.(c) Relation between corrected air flow and corrected engine speed.(d) Relation between corrected fuel consumption and corrected.engine speed.
(e) Relation between corrected specific fuel consumption based onnet thrust and corrected engine speed.
(f) Relation between corrected specific fuel consumption based onnet thrust and corrected net thrust.
(g) Relation between corrected fuel-air ratio and corrected enginespeed.
CONFIDENTIAL
-
20 CONFIDENTIAL NACA RM No. E7C13
Figure 9. - Effect of altitude on generalized performance character-iatics of 19XB-1 turbojet engine. Static test conditions. Per-formance parameters corrected to NACA standard atmospheric condi-tions at sea level.(a) Relation between corrected jet thrust and corrected engine
speed.(b) Relation between corrected air flow and corrected engine speed.(c) Relation between corrected fuel consumption and correctedengine speed,
(d) Relation between corrected specific fuel consumption based onjet thrust and corrected engine speed.
(e) Relation between corrected specific fuel consumption based onjet thrust and corrected jet thrust.
(f) Relation between corrected fuel-air ratio and corrected enginespeed.
Figure 10. - Effect of ram-pressure ratio on generalized performancecharacteristics of 19B-8 turbojet engine. Tail-pipe-nozzle area,106 square inches; simulated altitude, 15,000 feet. Performanceparameters corrected to NACA standard atmospheric conditions atsea level,(a) Relation between corrected jet thrust and corrected engine
speed.(b) Relation between corrected net thrust and corrected engine
speed.(c) Relation between corrected air flow and corrected engine speed.(d) Relation between corrected fuel consumption and correctedengine speed.
(e) Relation between corrected specific fuel consumption based onnet thrust and corrected engine speed.
(f) Relation between corrected specific fuel consumption based onnet thrust and corrected not thrust.
(L) Relation between specific fuel consumption based on net thrusthorsepower and corrected engine speed.
(h.) Relation between corrected. fuel•air ratio and corrected enginespeed.
Figure 11. - Effect of altitude on windmilling drag characteristicsof 19B-6 turbojet engine. Tail cone in.(a) Relation between corrected. windmilling drag and corrected. 'Grueairspeed, Windmilling drag and true airs-Deed corrected to NACAstandard atmospheric conditions at sea level.
(b) Relation between true airspeed and ratio of windmilling dragto maximum net thrust.
(c) Relation between corrected air flow and corrected engine wind-milling speed. Air flow and engine windmilling speed correctedto NACA standard atmospheric conditions at sea level.
CONFIDENTIAL
-
NACA RM NO. E7C13 CONFIDENTIAL Fig, I
(a) With cowling.
I b ) Without cowling,
Figure Is - installation of 19B turbojet engine in under—
slung nacelle for investigation in altitude wind tunnel,
CONFIDENTIAL
-
n
0
Mv
f^Oz'T1
e
OMz—a
Dr
nOz
OMz
Coa
Station 5
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Figure 2. — Longitudinal section of 19Bturbojet—engine installation showing measuringstations.
-n
iv
-
00
01®AO
a®
aa
000
0
ac^
.92
as^
064 ®90
7,0
NACA RM No. E7C13 CON FI DENT AL F g, 3
Corrected engine speed, N1416, rpm
Figure 3.- Relation between total-pressure ratio across cowl-inletscreen and corrected engine speed for range of simulatedaltitudes and flight speeds.
CONFIDENTIAL
-
36_
NACA RM NO. E703 CONFIDENTIAL Fig. 4a
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft)
O 5,000O 15,000
O 25,000
100
134
60
*' 40
a
20
I
6,000 8,000 10,000 12,000 14,000 16,000 18,EEngine speed, N, rpm
(a) Relation between jet thrust and engine speed.
Figure 4,- Effect of altitude on performance characteristics of198-8 turbojet engine. Tail-pipe-nozzle area, 135 squareinches; free -stream rare-pressure ratio, 1.013.
>00
CONFIDENTIAL
-
A
NACA RM NO. E7C13 CONFIDENTIAL Fi 9 . 4b
CONFIDENTIAL
-
1400
1200
1000
w
800
0
0 600Q
400
200
06 )00
NACA RPM No- E7C13 CONFIDENTIAL Fig. 4C
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, 135square inches; free — stream ram—pressure ratio, 1.013.
. CONFIDENTIAL
-
18,000
NACU, RM NO- E7 C i 3
CON FI DEN TI AL
Fig. 4d
r--
NATIONAL ADVISORYy COMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft)0 5,000p 15,000® 25,000
25
20
0Q)
15
e10
5
0
g-
-
5,
43a
s.
4.
4.
3.
43
4A
3,,
0100
2.
0
a.^
2Q
00H
U
U
P
NACA RAE Now E7Cf3 CON F! DENT AL F'g. 4e
00
CONFIDENTIAL
-
1018
.016
^aOw
A0
'w4 014
L
^-sOlt
®O1C
6, )00
NACA RM NO. E7CI3 CON FI DEBT! AL Fig. 4f
Eng ine speed, N, rpm(f) Relation between fuel-air ratio and engine speed.Pigwe 4.- Concluded. Effect of altitude on performance character-
istics of 13B-8 turbojet engine. Tail-pipe-nozzle area, 135square inches- free-stream ram-pressure ratio, 1,013.
CON FIDENTIAL
-
NACA RIB No. E7 13 CONFI DENTI AL Fi g. 5a
NATIONALONAL AObISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(f t)0 5,000q 10,000Q 15,000A 25,000V 30,000
3
aw
f,
a^^-a
400
01 I I m I I I I I I I I !6, 8,000 10,030 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
19 -1 turbojet engines Approximately static conditions.
CONFIDENTIAL
-
NACA Rte NO. E7C13 CONFIDENTIAL Fig. 5b
6-a
0
a
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft)
0 5,000q 10,000
400 ® 15,00025,000300000
209
000
800
600
400
200
05,000 8,000 10,000 12,000 14,000 16,000 189000
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 staticeonditionsm
CONFIDENTIAL
-
NACA RWI ,Now E7C13 CONFIDENTIAL F{g. 5c
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft)
0 5,000q 10,000
O 15800025,000
® 30¢000
25
20
a^cea
15
^i
ao 10
5
06,000 89000 10$000 12,000 14,000 16,000 188000
Engine speed, N, rpm
(c) Relation between air flow and engine speed,
Figure 6.— Continued. Effect of altitude on performance character -istics of 19XB--1 turbojet engine. Approximately static conditions.
CON F 1 DENT[ AL
-
N ACA RM No e E7 C 1 3 CON F I DEN TI AL F i 9 . 5d
S. 3.1
O
2.
cd
0
2.
0
Q)
6,000 8,000 10,000 12,000 148000 168000 18,000Engine speed, N, rpm
(d) Relation between specific fuel consumption based on jetthrust and engine speed,
Figure S.- Continued. Effect of altitude on performancecharacteristics of 19XB-1 turbojet enginee. Approximatelystatic conditions.
CONFIDENTIAL
-
NACA PM No. E7C13 CON FI DENT IAL Fig• 5e
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft )
0 5,000q 10,000O 15,000
25,000V 30,000
.018
.016
40
ro.014
f.
012
.010 L 1, 1 1 1 I I 1 1 1 I I 1,6 0001 5,000 10,000 120000 14,000 16,000 18,C
Engine speed, N, rpm
(e) 'Relation between fuel—air ratio and engine speed„
Figure 5.-- Concluded, Effect of altitude on performance charac-teristics of 19XB-1 turbojet engine„ Approximately staticconditions,
'vv
CONFIDENTIAL
-
NACA RM NO. E7CI3 CONFI DENTI AL Fi gn 6a
1000
800
600
400
200
0
CONFIDENTIAL
-
a^a
i
0
a^
(b) Relation between net thrust and engine speed®
Engine speed, W,
NACA RM NO. PC 1 3 CON FI DENTI AL ' Fig. 6b,, c
-
NACA RSA N0. E7C13 CONFI DENTI AL Fig. 6d
1200
1000
p-a
s
0
0v
800
60C
400
20C
C8,
Engine speed, N ,, rpm
CON FI DEN TI AE
-
NACA RM No. PC 13 CONFIDENTIAL Fig. 6e
a.'
A
4.0
3.5
Vr_0
Vii. 0
C0
0.2.5
U)V.00
^ 2.0
v
V)1.5
1.08,000 10,000 12,000 14,000 16,000 18,000
Engine speed, N, rpm
(e) Relation between specific fuel consumption based on netthrust and engine speed.
Figure 6.— Continued. Effect of free —stream ram—pressure ratioon performance characteristics of 1988 turbojet engine.Mail—pipe—nozzle area, 106 square inches; simulated altitude,15,000 feet.
CONFI DENTI AL
-
W N u . t I (,UNFI ULN I I AL Fe b
Of NATIONAL ADVISORY13.5
0COMMITTEE FOR AERONAUTICS
Free-streamram-
pressureratio
0 O 1.012O 1.122O 1.155
FN A C A
L-0
i
x.a^0P.w 6,0
in
5
a^
0
(D
0
P. 3.en
0v
o^
c-. 2
e^
v0
P. 1.
9..
0 ^_
3 I_
)I--
Li^ i sl3.000 10.000 12.000 14.000 16.000 18
Engine speed, N, rpm
(f) Relation between specific fuel consumption based on netthrust horsepower and engine speed.
Figure S.°- Continued. Effect of free-stream ram-pressure ratioon performance characteristics of 19B-8 turbojet engine.Tail-pipe-nozzle area, 106 square inches; simulated altitude,15,000 feet.
CON F I D E N T I A L
-
v
4.0
a^ 3.5
U)
s.
4j3,0
r_
0
2.5.o
z0
a.
2@0r0U
r-1
U1
1.5Uwi
V
Q
1.0
NACA RM NO. E7C13 CONFIDENTIAL Figs 6g
Net thrust, Fn v lb
(g) Relation between specific fuel consumption based on netthrust and net thrush
Figure 6, — Continued. Effect of free —stream ram—pressure ratioon performance characteristics of 19E-8 turbojet engine,Tall —pipe —nozzle area, 106 square Inches; simulated altitude„15,000 feet,
CONFIDENTIAL
-
NACA RIM N0, E7C13 CON FI DEN TI AL Fig, 6h
NATIONAL ADVISORYCOITTEE FOR AERONAUTICS
Free—streamram —
pressureratio
0 1,012q 1,122® 1.155
cd
0
a,
1020
1018
016
,014
012
.010
.008 .1 I I I I I I I I8,000 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—strews ram—pressure ratio
on performance characteristics of 198-8 turbojet engine.Tail —pipe—nozzle area, 106 square inches; simulated altitude,15,000 feet.
CON Ft DENT tAL
-
A
cn
4J
(Das
11
NACA RM No„ E7C13
CON FI DEN TI AL
Fig. 7a
I
gngine speed, N, rpm
(a) Relation between Jet thrust and engine speed.
Figure 7.- Effect of tail-pipe-nozzle area on performancecharacteristics of 19B-8 turbojet engine. Free-stream ram-pressure ratio, 1.131; simulated altitude, 10,000 feet.
CONFIDENTIAL
-
pF-4
^ 600
w
^^ 400
NACA RM ho. E7C13 CONFIDENTIAL Fig~ 7b
c
CONFIDENTIAL
-
1400
1200
1000
ra
800
r-0
600
00
d400
20C
NACA Rte No: E7C13 CONFI DENTI AL Fig, 7c
C
CONFIDENTIAL
-
NACA RIM NO. E7 C 1 3 CONFIDENTIAL Fi g . 7d
2!
2(
0, 1 11
8Engine speed, M, rpm
(d) Relation between air flow and engine speed.Figure 7.- Continued. Effect of tall-pipe-nozzle area on per-
formance characteristics of 199-8 turbojet engine. Free-streamram-pressure ratio, 1.131; simulated altitude, 10,000 feet.
CONFIDENTIAL
-
NACA RIM N®„ E7 13 CONFI DENTI AL Fi 9^ 7 e
a
LO
4P4
INI
4.
^4.0
r.0
a
.0/$
9 .0V
h^
gWq^^YYi
Y^Y
0ego
02.0
^gY
Engine speed, fib, rpm
(e) Relations between specific fuel consumption based on netthrust and engine speed.
Fi 7.m Continued. Effect of tail-pipe —nozzle area on per-tormance characteristics of 19B-6 turbojet engine. Free-scream
ram-pressure ratio, 1131; simulated altitude, 10,000 feet.CONFIDENTIAL
-
NACA RM Ig o„ E7C13 CON FI DEN TI AL Fig„ 7
a
7a0
0ca
6.0
0A
5. 0
^a
n0
4.0
10
0
5.0
00
2.0
0
c^ac^00 1N0
8 , 000 108000 12X000 14,000 168000 188000Engine speed, N, rpm
(f) Relation between specific fuel consumption based on netthrust horsepower and engine speed.
Figure 7„-. Continued. :Effect of tail-pipe —nozzle area on per-formance characteristics of 19B--8 turbojet en®inee Free-streamram-pressure ratio, 1,131; simulated altitude, 10,000 feet.
CONFIDENTIAL
-
5. 41
r
/
+S
P
^d
^^ PA
INI4d
4.1
ryry.A•5
y0g
^V,
Af./
0
k^
As.
r. ^e 4
A00
^4,
Aryl
P
V ^ P
(D
MJ
1.
NACA Rte NO. E7C13 CONFIDENTIAL Fig 7g
Net thrust, Fn, lb
(g) Relation between specific fuel consumption based on netthrust and net thrush
Figure 7b-- Continued. Effect of tail-pipe- . nozzle area on per-formance characteristics of 19B-8 turbojet engine. Free-screamram-pressure ratio, 1.131; simulated altitude, 10,000 feet.
CON F I DENT I A L
-
N ACA RM No, E7 C 13 CONFIDENTIAL F i 9 , 7n
.018
.016
e014rd
0
012
co
010
x.00E
. OOE6,000 10,000 12,000 14,000 16,000 18,000
Engine speed, N, rptA
(h) Relation between fuel-air ratio and engine speed.Figure 7 - Concluded, Effect of tail-pipe-nozzle area on per-
formance characteristics of 198-8 turbojet engine Free-streamram-pressure ratio, 1.131; simulated altitude, 10,000 feet
C ON F I D EN T I A L
-
H ACA RIB -No ^ E7 C 1 3 - COMFI DENTI AL Fi g,. 8a
NATIONAL ADVISORY
COMMMFE FOR AERONAUTICS
Simulatedaltitude
( f t)
0 5,0001400 q 15,000
,0 25,000
1200
1000
to1-1
800
s.
.600
4-)
40000
200
I I I __ I __L__ ___Jnoo io-onn 12 _nnn 14- non i r- - onn iR -noo 20
Corrected engine speed, W/F—.V V ' rpm(a) Relation between corrected jet thrust and corrected engine
speed,,Figure G.– Effect of altitude on generalized performance charac -
teristics of 195-8 turbojet engine, Free–stream ram –pressureratio, 1,013, Tail--pipe–nozzle area, 13.5 square inches,Performance parameters corrected to NACA - standard atmosphericconditions at sea level,
08, .0 000
CONFIDENTIAL
-
N A C A RU N o - E 7 C I 3 C O N F I D E N T I Ab F i g , 8 b
(b) Relation between c o ~ r e c t e d w e t t h ~ u s t a n d c o r r e c t e d engine speed,
F%gt.We Be- Cont%wued, E f i e c t of' af t f tu .de on genepelfeed per- f o ~ m a a e ekrrt~aetez-f sties of P9B-8 turbojet engf ne, Free- stmm ~ a p f 8 - p ~ s s ~ e p a t f o p 2 ,013 , Ta21-pipe-noz~le area, 3.35 squape fnches, P e ~ f o m a n c e p a ~ a m e t e r s c o ~ r e c t e d to H X C A standa~d abosphe r f e condi tfons at sea level,
C O N F I D E N T I A L
-
R A C A RbR Boo E s C 1 3 C O N F I DEN T I A L F i g . 8 c
Ffgure 8.- Continued. E f r e c t o r altitude on generallrad p'er- r o m a n c e charac t e r l s t l cs or 19B-8 t u r b o j e t e n g i n e . Free- s t ~ e a m ram-pressure r a t i o , 1.0l3. Tal l -p fpe -nozz le a r e a , 135 square i n c h e s , Perfol-mance parameters cors-ected t o NACA s t a n d a r d atrnospherf c eondf t i o n s a t sea l e v e l ,
-
1400
1200
C0
vaE
[ 10000U
80C
UNL,
0U
MACS RM No. E703 CONFI DENTI AL F+9. 8d
2000
1800
1600
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulated ' qqaltitude i
(ft)
0 5,000
0 1,5,000
O 25,000
/n
"Xi
60C
4008,000 10,000 12,000 14,000 16,000 18,000
Corrected engine speed, N14-9, rpm(d) Relation between corrected fuel consumption and corrected
engine speed,Figure 8,-° Continued, Effect of altitude on generalized per-
formance characteristics of 198°8 turbojet engine, Free–stream ram–pressure ratio,, 1„013 4 Tail–pipe–nozzle area,135 square inches. Performance parameters corrected to 'NACAstandard atmospheric conditions at sea level,
209000
CONFIDENTIAL
-
)00
No. E7C 1 3 CONFI DENTI AL Fig, 8eN AC A Rid
m.5
S»
. Ar-+
;01 5B0
G-.
4 5
.0
a^G
G0
ro3N a`
,O
C
O
a-9
3^C
0U
ri
2.U
6-^
UUSZ.
f!1
0 2.0vUUS.^a
O0 1
8,Corrected engine speed, N/fwg rpm
(e) Relation between corrected specific fuel consumption basedon net thrust and corrected engine speed,
Pigure 8. — Continued. Effect of altitude on generalized per-f°ormmnce characteristics of 19R-8 turbojet engine, Free—stream ram —pressure ratio, 1,013, 'Paid.—pipe—nozzle area,135 square inches, Performance parameters corrected to i'dACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
5.0
4.5
ft
410
010
3.5
ro
r.m
3.®
005-a
$14 2,5
v^s
2ryOwa'0
0VI
1.5
Corrected net thrust, Fn/Fa, lb
NACA R ye NO. E7C13 CONFIDENTIAL Fig 8f
(f) Relation between corrected specific fuel consumption based onnet thrust and corrected net thrust.
Figure 8.- Continued, Effect of altitude on generalized performancecharacteristics of 198-5 turbojet engine. Free-stream ram-pressureratio, 1.013. Tail-pipe-nozzle area, 135 square inches, Perform -ance parameters corrected to NACA standard atmospheric conditionsat sea: l.evel^
CON F I DFN T I A L
-
9
)00
ft0
c^
^s0
o
c^
0U
10
NACA RM H0. E7C13 CONFIDENTIAL Fig - 89
Corrected engine speed, N/fe—, rpm(g) Relation between corrected fuel—air.ratio and corrected
engine speed.
Fi^ re S.— Concluded. Effect.of altitude on generalized per—for -ante characteristics of 19B-8 turbojet engine, Free—stream ram —pressure ratio, 1.013, Tail —pipe—nozzle area,135 square inches. Performance parameters corrected to NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
V4)a^0
00
NACA Rho No. E7 C1 3 CONFIDENTIAL F+9e 9a
Corrected engine speed, fit/J9, rpmW. `gelation between oorrected jet thrust and corrected enginespeed,Pigure 9.- Effect of altitude on generalized performance charac-
teristics of 19XB-1 turbojet engine. Static test conditions.Perrorma ce parameters corrected to WACA standard atmosphericconditions at sea. level,
CONFIDENTIAL
-
^
tcofis^
BPI0H^A
c^
va^v
00
NACA Rte loo E7CI3 CON FI DENT IAL Fig 9b
30
25
20
15
10
5
C"000 QP000 110 000 13,000 is P 000 17,,000
Corrected engine speed, N/F, rpm(b) Relation between corrected air flora, and corrected engine
speed.Figure 9,— Continued. Effect of altitude on generalized per-
formance characteristics of 19XB-1 turbojet engine. Statictest conditions. Performance parameters corrected to NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
f€ ACA RM Noe E7C13
CONFI DENTI AL Fig. 9c
2000
1000
1600
1.400
01200
00
aH 1000
^a
S0000
600
400
^.,
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
Simulatedaltitude
(ft)
0 5,0000 10,000(> 15"000A 25,000 V 30$000
lu
77' 000 99000 11"000 13,000 15"000 17P000Corrected engine speed, /F, rpm
(c) Relation between corrected fuel, consumption and correctedengine speed.
Figure 9.-a Continued,, Effect of altitude on generalized per-formance characteristics of 19X5-1 turbojet engine. Statictest conditions, Performance parameters corrected to NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
.s^
0
E4
c..
C=-
a.'
z.
c0
`0 3a^
p
Oa^a.,
2E^nc0U
r-i
c, 2 .1
V
4—r
+ra
UQ7
Ql.0U
N
0 1U
NACA RM NO. E7 1 3
CONFIDENTIAL Fi g^ 9d
7,000 9,000 110000 13,000 15,000 17,000Corrected engine speed, N/ NW, rprr.
(d) Relation between corrected specific fuel consumption basedon jet thrust and corrected engine speed,
Figure 9.— Continued, Effect of altitude on generalized peer°formance characteristics of 19XB-1 turbojet engine. Statictest conditions, Performance parameters corrected to NACA.standard atmospheric conditions at sea level,
CONFIDENTIAL
-
NACA RJR No. E7C13
C 0 N F I DEN TI AL Fi g t 9e
a
3.5
4Jn^
0
3.0
Cd
0
2.5
tor.0U
a^
c-, 2.0
UK-1
4a
ari
U
U
Q
1 ® ^
v
UUF+
F+OU 1.0
ICorrected ,jet thrust, F,/6, lb
(e) Relation between corrected specific fuel consumption based onJet thrust and corrected jet thrust.
Figure 9.— Continued. Effect of altitude on generalized performancecharacteristics of 19XB-1 turbojet engine. Static test conditions,Performance parameters corrected to NACA standard atmosphericconditions at sea level.
CONFIDENTIAL
-
NACA RNI NO. E703
CON FI DENT AL
Fi g. 9 f
.020
W .018
ro
0
a .016roF-.
ro
.014
0.sue0c^
s, 01200
.0107.
Corrected engine speed, N14T, 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, Statictest conditions. Performance parameters corrected to NACAstandard atmospheric conditions at sea level.
CONFI DENTIAL
-
NACA RM NO. E 7 C13 CON FI DEN TI AL Fig. 10a
18(
Free—streamram—
16(
pressureratio
O 1 0121122
O 1 ry 155
14(
121 )0
w
10
c.
QJ
p"^ S
TiNdJ
UN
OU
6
4
280 1-O 10,000 12,000 14,000 16,000 1$,000
Corrected engine speed, N/44, rpm(a) Relation between cor&rect'ed jet thrust and corrected engine
speed,Figure 10.- Effect of ram-pressure ratio on generalized per-
formance characteristics of 19R-8 turbojet engine ¢ Tail-pipe—nozzle area, 106 square inches; simulated altitude,15,000 feet, Performance parameters corrected to NACAstandard atmospheric conditions at sea level.
DO
CONFIDENTIAL
-
AH
.s^
43
0
00
NACA RM NO- E7C13 CON FI DEN TIAL Fig. 10b
8V000 10,000 12,000 14,000 16,000 189000Corrected engine speed, N/-A_), rpm
(b) Relation between corrected net thrust and corrected enginespeed,
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,16,000 feet, Performance parameters corrected to NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
NACA RMt No. E7C13 CON FI DENT AL Fig. 10e
35
30
5
^0
0H^a
cd
Vaa'e^^a
0U
5
08,
Corrected engine speed, N/^95, rpm(c) Relation between corrected air flow and corrected engine
speed.
Figure 10. — Continued. Effect of ram. —pressure ratio on gener-alized performance characteristics of 198-8 turbojet engine.'fail—pipe—nozzle area, 106 square inches; simulated altitude,15,000 feet. Performance parameters corrected to HACAstandard atmospheric conditions at sea level,
20
15
10
CONFIDENTIAL
-
F,
rl
0
a
0v
os
aa,
U
$rOU
NACA Reel NO. E7C13 CON FI DEN TI AL Fig. 10d
(d) Relation between corrected fuel consumption and correctedengine 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 NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
ve
Av
^e 4.v
s. 4
^ g
^g
3 .
^e
N
0 3a
b0U)
0 2.
Q.E
0
v.^ 2 ry
v
a.
U5
dU
NACA RJR No. E7C13 CONFIDENTIAL F+g. 10
0 8,000 10,000 12,000 14,000 16,000 18,000UCorrected engine speed, Fd14-6 $ rpm
(e) Relation between corrected specific fuel consumption basedon net thrust and corrected engine speed,,
Figure 10.- Continued, Effect of ram-pressure ratio on gener-alized performance characteristics of 198-8 turbolet engine,Tail-pipe nozzle area, 106 squar=e inches; simulated altitude,16 & 000 feet, Performance parameters corrected to WACAstandard atmospheric conditions at sea level„
CONFIDENTIAL
-
NACA RM NO. E7C13 CON FI DENT IAL Fig, 10
aA
4g5
4.0
4J
43545
4J4)r.Z0
3.00
Z0
41
00U
@a 2„0
U
ie
U
1.5
0)
o 1.0v 200 400 600 800 1000 1200 lane
Corrected net thrust, Pn/6, lb(f) Relation between corrected specific fuel consumption based on net
thrust and corrected net thrust.
Figure 10. — Continued. Effect of rare—pressure ratio on generalizedperformance characteristics of 19B-8 turbojet engine. Tail—pipe -nozzle area, 106 square inches; simulated altitude, 15,000 feet,Performance parameters corrected to NACA standard atmosphericconditions at sea level,
CONFIDENTIAL
-
900
cl,
^a
0790
0aanti0'G
6.04J
a
A`a 8.0
Free--streamram —
pressureratio
O 1@0120 1.122p 1.155
NACA Rte No. E7C13 CON FI DEN TI AL Fig. 10g
NATIONAL ADVISORYCO WEE FOR AERONAUTICS
C: 5.0c0bU)Cd
400
0
4-)o.a
0 5e00
a^
v2,0
UaLC2.
tf^
1.08
10 MC) 19. Mr) 1A nr*ri i^ nnn ^A nnn
Corrected engine speed, Nl -^6 6 g rpm --n(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 188°8 turbojet engine.Tail—pipe —nozzle area, 106 square inches; simulated altitude,,15,000 feet, Performance parameters corrected to WACAstandard atmospheric conditions at sea level,
CONFI DENTI AL
-
NACA RM NO. E7 C 1 3 CONPI DENTIAL F i g w 10h
.022
.020
.018
Cd
0
016as
014
0120
U
.010
008a
Corrected engine speed, Nl q-o ,, rpm(h) Relation between corrected fuel—air ratio and corrected
engine speed,
Figure 10. — Concluded. Effect of ram—pressure ratio on gener-alized performance characteristics of 198•8 turbojet engine,Tail—pipe—nozzle area, 106 square inches-, simulated altitude,,15,000 feet, Performance parameters corrected to NACAstandard atmospheric conditions at sea level,
CONFIDENTIAL
-
N ACA RM No. E7C 1 3
CON F I DEN T I AL
F i 9 . I I a
NATIONAL ADVISORYCOMMITTEE FOP AERONAUTICS
)o I I I
11
11
10rl
to
cd
bo
rM
00
sow
2M-
-
'SD u2
05
,15
IN zSimulatedaltitude
(ft )
0 5,000
0 20,0000 30,000
N AC A RM N o . E7 C 1 3
Fi 9 ° I I b, c
NATIONAL ADVISORYCOMMITTEE FOR AERONAUTICS
0 0 100 j 200 300 400 500True airspeed, W, mph0
(b) Relation between true airspeed and ratio of windmilling dragto maximum net thrust®
15
hF
to
3C 10
0ifs
0rrhd
5
0DO
Corrected engine windmilli .ng speed, N14W, rpm(c) Relation between corrected air flow and corrected engine wind—
milling speed, Air flow and engine windmilling speed correctedto WAGA, standard atmospheric conditions at sea level,
Figure I.I.— Concluded, Effect of altitude on windmilling dragcharacteri sties of 19B--8 turbojet engine, Tall—pipe^no7zle ares.,135 square inches,
Restriction/Classification Cancelled
Restriction/Classification Cancelled