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  • Aircraft Propulsion and Power  (37)
  • 1950-1954  (37)
  • 1
    Publication Date: 2014-07-15
    Description: A literature survey was conducted to determine the relation between aircraft ignition sources and inflammables. Available literature applicable to the problem of aircraft fire hazards is analyzed and, discussed herein. Data pertaining to the effect of many variables on ignition temperatures, minimum ignition pressures, and minimum spark-ignition energies of inflammables, quenching distances of electrode configurations, and size of openings incapable of flame propagation are presented and discussed. The ignition temperatures and the limits of inflammability of gasoline in air in different test environments, and the minimum ignition pressure and the minimum size of openings for flame propagation of gasoline - air mixtures are included. Inerting of gasoline - air mixtures is discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TN-2227
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  • 2
    Publication Date: 2014-07-15
    Description: An experimental investigation was conducted to determine the cooling effectiveness of a wide variety of air-cooled turbine-blade configurations. The blades, which were tested in the turbine of a - commercial turbojet engine that was modified for this investigation by replacing two of the original blades with air-cooled blades located diametrically opposite each other, are untwisted, have no aerodynamic taper, and have essentially the same external profile. The cooling-passage configuration is different for each blade, however. The fabrication procedures were varied and often unique. The blades were fabricated using methods most suitable for obtaining a small number of blades for use in the cooling investigations and therefore not all the fabrication procedures would be directly applicable to production processes, although some of the ideas and steps might be useful. Blade shells were obtained by both casting and forming. The cast shells were either welded to the blade base or cast integrally with the base. The formed shells were attached to the base by a brazing and two welding methods. Additional surface area was supplied in the coolant passages by the addition of fins or tubes that were S-brazed. to the shell. A number of blades with special leading- and trailing-edge designs that provided added cooling to these areas were fabricated. The cooling effectiveness and purposes of the various blade configurations are discussed briefly.
    Keywords: Aircraft Propulsion and Power
    Type: REPT-2203 , NACA-RM-E51E23
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  • 3
    Publication Date: 2014-07-15
    Description: Sound pressure levels, frequency spectrum, and jet velocity profiles are presented for an engine-afterburner combination at various values of afterburner fuel - air ratio. At the high fuel-air ratios, severe low-frequency resonance was encountered which represented more than half the total energy in the sound spectrum. At similar thrust conditions, lower sound pressure levels were obtained from a current fighter air craft with a different afterburner configuration. The lower sound pressure levels are attributed to resonance-free afterburner operation and thereby indicate the importance of acoustic considerations in afterburner design.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E54G07
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  • 4
    Publication Date: 2014-07-15
    Description: Tests of two propellers having two blades and differing only in the inboard pitch distribution were made in the Langley 8-foot highspeed tunnel to determine the effect of inboard pitch distribution on propeller performance. propeller was designed for operation in the reduced velocity region ahead of an NACA cowling; the inboard pitch distribution of the modified propeller was increased for operation at or near free-stream velocities, such as would be obtained in a pusher installation. conditions covering climb, cruise, and high-speed operation. Wake surveys were taken behind the propellers in order to determine the distribution of thrust along the blades and to aid in the analysis of the results. Test results showed that the modified propeller was about 2.5 percent less efficient for a typical climb condition at all altitudes, 2 percent more efficient for one cruise condition, and 5 percent more efficient for high-speed operation. speed condition, the modified propeller showed a 6-percent loss in efficiency due to compressibility; whereas the original propeller showed an 11-percent efficiency loss due to compressiblity. The lower compressibility loss for the modified propeller resulted from the fact that the inboard sections of this propeller could operate at increased thrust loading after compressibility losses had occurred at the outboard sections.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TN-2268
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  • 5
    Publication Date: 2019-06-28
    Description: A comparison of the operating characteristics of 75-millimeter-bore (size 215) cylindrical-roller one-piece inner-race-riding cage-type bearings was made using a laboratory test rig and a turbojet engine. Cooling correlation parameters were determined by means of dimensional analysis, and the generalized results for both the inner- and outer-race bearing operating temperatures are compared for the laboratory test rig and the turbojet engine. Inner- and outer-race cooling-correlation curves were obtained for the turbojet-engine turbine-roller bearing with the same inner- and outer-race correlation parameters and exponents as those determined for the laboratory test-rig bearing. The inner- and outer-race turbine roller-bearing temperatures may be predicted from a single curve, regardless of variations in speed, load, oil flow, oil inlet temperature, oil inlet viscosity, oil-jet diameter or any combination of these parameters. The turbojet-engine turbine-roller-bearing inner-race temperatures were 30 to 60 F greater than the outer-race-maximum temperatures, the exact values depending on the operating condition and oil viscosity; these results are in contrast to the laboratory test-rig results where the inner-race temperatures were less than the outer-race-maximum temperatures. The turbojet-engine turbine-roller bearing, maximum outer-race circumferential temperature variation was approximately 30 F for each of the oils used. The effect of oil viscosity on inner- and outer-race turbojet-engine turbine-roller-bearing temperatures was found to be significant. With the lower viscosity oil (6x10(exp -7) reyns (4.9 centistokes) at 100 F; viscosity index, 83), the inner-race temperature was approximately 30 to 35 F less than with the higher viscosity oil (53x10(exp -7) reyns (42.8 centistokes) at 100 F; viscosity index, 150); whereas the outer-race-maximum temperatures were 12 to 28 F lower with the lower viscosity oil over the DN range investigated.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51I05
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  • 6
    Publication Date: 2019-07-12
    Description: A theoretical method for evaluating the stability characteristics and the amplitude and the frequency of pulsation of ram-jet engines without heat addition is presented herein. Experimental verification of the theoretical results are included where data were available. Theory and experiment show that the pulsation amplitude of a high mass-flow-ratio diffuser having no cone surface flow separation increases with decreasing mass flow. The theoretical trends for changes in amplitude, frequency, and mean-pressure recovery with changes in plenum-chamber volume were experimentally confirmed. For perforated convergent-divergent-type diffusers, a stability hysteresis loop was predicted on the pressure-recovery mass-flow-ratio curve. At a given mean mass-flow ratio, the higher.value of mean pressure recovery corresponded to oscillatory flow in the diffuser while the lower branch was stable. This hysteresis has been observed experimentally. The theory indicates that for a ram-jet engine of given diameter, the amplitude of pulsation of a supersonic diffuser is increased by decreasing the relative size of the plenum chamber with respect to the diffuser volume down to a critical value at which oscillations cease. In the region of these critical values, the stable mass-flow range of the diffuser may be increased either by decreasing the combustion chamber volume or by increasing the length of the diffuser.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E52I24
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  • 7
    Publication Date: 2019-07-11
    Description: A method has been developed for modifying a rocket motor so that its exhaust characteristics simulate those of a turbojet engine. The analysis necessary to the design is presented along with tests from which the designs are evaluated. Simulation was found to be best if the exhaust characteristics to be duplicated were those of a turbojet engine at high altitudes and with the afterburner operative.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-L54I15
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  • 8
    Publication Date: 2019-07-11
    Description: An investigation was conducted at simulated high-altitude flight conditions to evaluate the use of compressor evaporative cooling as a means of turbojet-engine thrust augmentation. Comparison of the performance of the engine with water-alcohol injection at the compressor inlet, at the sixth stage of the compressor, and at the sixth and ninth stages was made. From consideration of the thrust increases achieved, the interstage injection of the coolant was considered more desirable preferred over the combined sixth- and ninth-stage injection because of its relative simplicity. A maximum augmented net-thrust ratio of 1.106 and a maximum augmented jet-thrust ratio of 1.062 were obtained at an augmented liquid ratio of 2.98 and an engine-inlet temperature of 80 F. At lower inlet temperatures (-40 to 40 F), the maximum augmented net-thrust ratios ranged from 1.040 to 1.076 and the maximum augmented jet-thrust ratios ranged from 1.027 to 1.048, depending upon the inlet temperature. The relatively small increase in performance at the lower inlet-air temperatures can be partially attributed to the inadequate evaporation of the water-alcohol mixture, but the more significant limitation was believed to be caused by the negative influence of the liquid coolant on engine- component performance. In general, it is concluded that the effectiveness of the injection of a coolant into the compressor as a means of thrust augmentation is considerably influenced by the design characteristics of the components of the engine being used.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E52F20
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  • 9
    Publication Date: 2019-07-11
    Description: Performance data obtained with recording oscillographs are presented to show the transient response of the General Electric Integrated Electronic Control operating on the J47 RXl-3 turbo-Jet engine over a range of altitudes from 10,000 to 45,000 feet and at ram pressure ratios of 1.03 and 1.4. These data represent the performance of the final control configuration developed after an investigation of the engine transient behavior in the NACA altitude wind tunnel. Oscillograph traces of controlled accelerations (throttle bursts),oontrolled decelerations (throttle chops), and controlled altitude starts are presented.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50G12
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  • 10
    Publication Date: 2019-07-11
    Description: A modified J33-A-27 compressor was operated over a range of equivalent impeller speeds from 6100 to 13,250 rpm in order to obtain the over-all compressor performance. At the equivalent design speed of 11,800 rpm, the maximum efficiency of 0.764 and peak pressure ratio of 4.56 occurred at an equivalent weight flow of 104.07 pounds per second. At the highest equivalent speed (13,250 rpm) a maximum efficiency of 0.711, a maximum equivalent weight flow of 123.00 pounds per second, and a peak pressure ratio of 5.76 were obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50D25
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  • 11
    Publication Date: 2019-07-12
    Description: At the request of the Bureau of Aeronautics, Department of the Navy, an investigation of the Westinghouse XJ34-WE-32 turbojet engine is being conducted in the NACA Lewis altitude wind tunnel to determine the steady-state and transient operating characteristics of the controlled and uncontrolled engine at various altitudes and ram pressure ratios. As part of this program, transient performance data that illustrate the operation of the engine is obtained in the form of oscillographic traces. Similar data for engine operation i n the afterburning range, covering a range of throttle settings from the minimum value giving rated speed (throttle position, 72 degrees) to full afterburning (throttle position, ll0 degrees), is presented herein. These data thus serve to indicate the transient characteristics of the engine when the throttle is advance into, withdrawn from, and moved within the afterburning range in a stepwise manner, as well as the steady-state stability of the engine during afterburning .
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50L29
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  • 12
    Publication Date: 2019-07-12
    Description: The operational characteristics of a J57-P1 turbojet engine have been investigated at altitudes between 15,000 and 66,000 feet in the Lewis altitude wind tunnel. Included in this study is a discussion of fuel nozzle coking, the altitude operating limits with and without the standard engine control, the compressor surge characteristics, and the engine starting and windmilling characteristics. Severe circumferential turbine outlet temperature gradients which occurred at high altitude as a result of fuel nozzle coking were alleviated by the manufacturer's change in the fuel flow divider schedule and in a nozzle gasket material. Compressor air bleed is required to prevent surge of the outboard compressor in the low engine speed region. The maximum altitude at which the engine was operated without the control was about 66,000 feet at 0.8 flight Mach number and at a reduced engine speed to avoid compressor surge; with the engine control in operation, the altitude operating limit is reduced to approximately 59,000 feet. The maximum altitude at which the engine was started was about 40,000 feet.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE54C31
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  • 13
    Publication Date: 2019-07-12
    Description: The stator-blade angles in the twelfth to fifteenth stages of a 16-stage high-pressure-ratio axial-flow compressor were decreased 3 deg The over-all performance of this compressor is compared with the performance of the same compressor with standard blade angles. The matching characteristics of the modified compressor and a two-stage turbine were also obtained and compared with those of the compressor with the original blade angles and the same turbine.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51L03
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  • 14
    Publication Date: 2019-07-12
    Description: An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations. The station near the entrance section of the combustors indicated a peak pressure ratio of 6.3 at an adiabatic efficiency of 0.63 for a corrected weight flow of 23.1 pounds per second. The other, located one blade-chord downstream of the last stator row, indicated a peak pressure ratio of 6.97 at an adiabatic efficiency of 0.81 for a corrected weight flow of 30.4 pounds per second. The difference in performance obtained at the two stations is attributed to shock waves in the vicinity of the last stator row. These shock waves and the accompanying flow choking, together with interstage circulatory flows, shift the compressor operating curves into the region where surge would normally occur. The inability of the compressor to meet design pressure ratio is probably due to boundary-layer buildup in the last stages, which cause axial velocities greater than design values that, in turn, adversely affect the angles of attack and turning angles in these blade rows.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50E22
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  • 15
    Publication Date: 2019-07-12
    Description: An investigation was conducted at the NACA Lewis laboratory to determine whether simulated porous gas-turbine blades fabricated by the Eaton Manufacturing Company of Cleveland, Ohio would be satisfactory with respect to coolant flow for application in gas-turbine engines. These blades simulated porous turbine blades by forcing the cooling air onto the blade surface through a large number of chordwise openings or slits between laminations of sheet metal or wire. This type of surface has a finite number of openings, whereas a porous surface has an almost infinite number of smaller openings for the coolant flow. The investigation showed that a blade made of sheet-metal laminations stacked on a support member that passed up through the coolant passage was completely unsatisfactory because of extremely poor coolant flow distribution over the blade surface. The flow distribution for two wire-wound blades was more uniform, but the pressure drop between the coolant supply pressure and the local pressure on the outside of the blades was too low by a factor ranging from 3 to 3.5 for the required coolant flow rates. The pressure drop could be increased by forcing the wires closer together during blade fabrication.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE51C13
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  • 16
    Publication Date: 2019-07-12
    Description: An investigation to increase the compressor surge-limit pressure ratio of the XJ40-WE-6 turbojet engine at high equivalent speeds was conducted at the NACA Lewis altitude wind tunnel. This report evaluates the compressor modifications which were restricted to (1) twisting rotor blades (in place) to change blade section angles and (2) inserting new stator diaphragms with different blade angles. Such configuration changes could be incorporated quickly and easily in existing engines at overhaul depots. It was found that slight improvements in the compressor surge limit were possible by compressor blade adjustment. However, some of the modifications also reduced the engine air flow and hence penalized the thrust. The use of a mixer assembly at the compressor outlet improved the surge limit with no appreciable thrust penalty.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE52G03
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  • 17
    Publication Date: 2019-07-12
    Description: A program was undertaken to determine the J73 turbojet engine compressor stall and surge characteristics and combustor blow-out limits enc ountered during transient engine operation. Data were obtained in the form of oscillograph traces showing the time history of several engi ne parameters with changes in engine fuel flow. The data presented in this report are for step and ramp changes in fuel flow at an altitude of 45,000 feet and flight Mach numbers of 0 and 0.8.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE53F30
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  • 18
    Publication Date: 2019-07-12
    Description: A turbine blade with a porous stainless-steel shell sintered to a supporting steel strut has been fabricated for tests at the NACA by Federal-Mogul Corporation under contract from the Bureau of Aeronautics, Department of the Navy. The apparent permeability of this blade, on the average, more nearly approaches the values specified by the NAGA than did two strut-supported bronze blades in a previous investigation. Random variations of permeability in the present blade are substantialy greater than those of the bronze blades, but projected improvements in certain phases of the fabrication process are expected to reduce these variations.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE54D29
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  • 19
    Publication Date: 2019-07-12
    Description: The performance of a 13-stage development comressor for the J40-WE-24 engine has been determined at equivalent speeds from 30 to 112 percent of design. The design total-pressure ratio of 6.0 and the design weight flow of 164 pounds per second were not attained, An analysis was conducted to determine the reasons for the poor performance at the design and over-design speed. The analysis indicated that most of the difficulty could be attributed to the fact that the first stage was overcompromised to favor part-speed performance,
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE53D17
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  • 20
    Publication Date: 2019-07-12
    Description: This report summarizes the effects of fuel volatility and engine design variables on the problem of starting gas-turbine engines at sea-level and altitude conditions. The starting operation for engines with tubular combustors is considered as three steps; namely, (1) ignition of a fuel-air mixture in the combustor, (2) propagation of flame through cross-fire tubes to all combustors, and (3) acceleration of the engine from windmilling or starting speed to the operating speed range. Pertinent data from laboratory researches, single-combustor studies, and full-scale engine investigations are presented on each phase of the starting problem.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE51B02
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  • 21
    Publication Date: 2019-07-12
    Description: Altitude performance of a YJ71-A-7 turbojet engine, with afterburner inoperative, was determined in the NACA Lewis altitude wind tunnel over a wide range of flight conditions. Engine speed and exhaust-nozzle area were controlled independently during this investigation. The variation of corrected values of air flow, net thrust, and fuel flow with corrected engine speed was not defined by a single curve with changes in altitude at given flight Mach number. Changes in altitude had very little effect on minimum specific fuel consumption at altitudes up to 45,000 feet. There is one exhaust-nozzle schedule that is nearly optimum for all flight conditions. Performance calculated from pumping characteristics agreed with experimental values and can therefore be used to extend engine performance data.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E53E13
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  • 22
    Publication Date: 2019-07-12
    Description: A .General Electric fuel and torque regulator was tested in conjunction with a T31-3 turbine-propeller engine in the sea-level static test stand at the NACA Lewis laboratory. The engine and control were operated over the entire speed range: 11,000 rpm, nominal flight idle, to 13,000 rpm, full power. Steady-state and transient data were recorded and are presented with a description of the four control loops being used in the system. Results of this investigation indicated that single-lever control operation was satisfactory under conditions of test. Transient data presented showed that turbine-outlet temperature did overshoot maximum operating value on acceleration but that the time duration of overshoot did not exceed approximately 1 second. This temperature limiting resulted from a control on fuel flow as a function of engine speed. Speed and torque first reached their desired values 0.4 second from the time of change in power-setting lever position. Maximum speed overshoot was 3 percent.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE1H20
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  • 23
    Publication Date: 2019-07-11
    Description: Strain-gages were used to measure blade vibrations causing failures in the third stage of a production 11-stage axial-flow compressor. After the serious third-stage vibration was detected, a series of investigations were conducted with second-stage vane assemblies of varying angles of incidence. Curves presented herein show the effect of varying the angle of incidence of second-stage vane assembly on third-stage rotor-blade vibration amplitude and engine performance. A minimum vibration amplitude was obtained without greatly affecting the engine performance with a second-stage vane assembly of 9deg. greater angle of incidence than the assembly normally furnished with the engine.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE51F08
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  • 24
    Publication Date: 2019-07-11
    Description: The compressor from the XT-46 turbine-propeller engine was revised by removing the last two rows of stator blades and by eliminating the interstage leakage paths described in a previous report. With the revised compressor, the flow choking point shifted upstream into the last rotor-blade row but the maximum weight flow was not increased over that of the original compressor. The flow range of the revised compressor was reduced to about two-thirds that obtained with the original compressor. The later stages of the compressor did not produce the design static-pressure increase probably because of excessive boundary-layer build-up in this region. Measurements obtained in the ninth-stage stator showed that the performance up to this station was promising but that the last three stages of the compressor were limiting the useful operating range of the preceding stages. Some modifications in flow-passage geometry and blade settings are believed to be necessary, however, before any major improvements in over-all compressor performance can be obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50J10
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  • 25
    Publication Date: 2019-07-11
    Description: An investigation was conducted to determine the effects of water injection on the over-all performance of a modified J33-A-27 turbojet-engine compressor at the design equivalent speed of 11,800 rpm. The water-air ratio by weight was 0.05. With water injection the peak pressure ratio increased 9.0 per- cent, the maximum efficiency decreased 15 percent (actual numerical difference 0.12), and. the maximum total weight flow increased 9.3 percent.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50F14
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  • 26
    Publication Date: 2019-07-11
    Description: An investigation of a decoupler and a controlled-feathering device incorporated with the YT-56A turboprop engine has been made to determine the effectiveness of these devices in reducing the high negative thrust (drag) which accompanies power failure of this type of engine. Power failures were simulated by fuel cut-off, both without either device free to operate, and with each device free to operate singly. The investigation was made through an airspeed range from 50 to 230 mph. It was found that with neither device free to operate, the drag levels realized after power failures at airspeeds above 170 mph would impose vertical tail loads higher than those allowable for the YC-130, the airplane for which the test power package was designed. These levels were reached in approximately one second. The maximum drag realized after power failure was not appreciably altered by the use of the decoupler although the decoupler did put a limit on the duration of the peak drag. The controlled-feathering device maintained a level of essentially zero drag after power failure. The use of the decoupler in the YT-56A engine complicates windmilling air-starting procedures and makes it necessary to place operating restrictions on the engine to assure safe flight at low-power conditions,
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SA54I09
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  • 27
    Publication Date: 2019-07-12
    Description: A program was undertaken to determine the J73 turbojet engine compressor stall and surge characteristics and combustor blow-out limits encountered during transient engine operation. Data were obtained in the form of oscillograph traces showing the time history of several engine performance parameters with changes in engine fuel flow. The data presented in this report are for step changes in fuel flow at an altitude of 35,000 feet, at flight Mach numbers of 0.3, 0.8, and 1.2, and at several engine-inlet temperatures,
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE53F29
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  • 28
    Publication Date: 2019-07-12
    Description: Component data on the J35-A-23 compressor and two-stage turbine were used to determine the problems in matching the two units for operatio n in a turbojet engine. Possible operating regions were determined an d an equilibrium operating line was also determined for the assumed c onditions of zero flight speed and a jet nozzle area approximately 5. 5 percent greater than the wide-open nozzle area.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51H15
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  • 29
    Publication Date: 2019-07-12
    Description: An investigation of the effect of inlet pressure, corrected engine speed, and turbine temperature level on turbine-inlet gas temperature distributions was conducted on a J40-WE-6, interim J40-WE-6, and prototype J40-WE-8 turbojet engine in the altitude wind tunnel at the NAC.4 Lewis laboratory. The engines were investigated over a range of simulated pressure altitudes from 15,000 to 55,000 feet, flight Mach numbers from 0.12 to 0.64, and corrected engine speeds from 7198 to 8026 rpm, The gas temperature distribution at the turbine of the three engines over the range of operating conditions investigated was considered satisfactory from the standpoint of desired temperature distribution with one exception - the distribution for the J40-WE-6 engine indicated a trend with decreasing engine-inlet pressure for the temperature to exceed the desired in the region of the blade hub. Installation of a compressor-outlet mixer vane assembly remedied this undesirable temperature distribution, The experimental data have shown that turbine-inlet temperature distributions are influenced in the expected manner by changes in compressor-outlet pressure or mass-flow distribution and by changes in combustor hole-area distribution. The similarity between turbine-inlet and turbine-outlet temperature distribution indicated only a small shift in temperature distribution imposed by the turbine rotors. The attainable jet thrusts of the three engines were influenced in different degrees and directions by changes in temperature distributions with change in engine-inlet pressure. Inability to match the desired temperature distribution resulted, for the J40-WE-6 engine, in an 11-percent thrust loss based on an average turbine-inlet temperature of 1500 F at an engine-inlet pressure of 500 pounds per square foot absolute. Departure from the desired temperature distribution in the Slade tip region results, for the prototype J40-WE-8 engine, in an attainable thrust increase of 3 to 4 percent as compared with that obtained if tip-region temperature limitations were observed.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E52H06
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  • 30
    Publication Date: 2019-07-11
    Description: An investigation of the altitude performance characteristics of an Allison J35-A-17 turbojet engines have been conducted in an altitude chamber at the NACA Lewis laboratory. Engine performance was obtained over a range of altitudes from 20,000 to 60,000 feet at a flight Mach number of 0.62 and a range of flight Mach numbers from 0.42 to 1.22 at an altitude of 30,000 feet. The performance of the engine over the range investigated could be generalized up to an altitude of 30,000 feet. Performance of the engine at any flight Mach number in the range investigated can be predicted for those operating condition a t which critical flow exits in the exhaust nozzle with the exception of the variables corrected net thrust, and net-thrust specific fuel consumption.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E50I15
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  • 31
    Publication Date: 2019-07-11
    Description: The power plant from a Mark 25 aerial torpedo was investigated both as a two-stage turbine and as a single-stage modified turbine to determine the effect on overall performance of nozzle size and shape, first-stage rotor-blade configuration, and axial nozzle-rotor running clearance. Performance was evaluated in terms of brake, rotor, and blade efficiencies. All the performance data were obtained for inlet total to outlet static pressure ratios of 8, 15 (design), and 20 with inlet conditions maintained constant at 95 pounds per square inch gage and 1000 F for rotor speeds from approximately 6000 to 18,000 rpm.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50D12
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  • 32
    Publication Date: 2019-07-11
    Description: This report presents a compilation of static sea-level data on existing or designed American and British axial-flow turbojet engines in terms of basic engine parameters such as thrust and air flow. In the data presented, changes in the over-U engine performance with time sre examined as well as the relation of the various engine parameters to each other.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-51K29
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  • 33
    Publication Date: 2019-07-10
    Description: An investigation was made of the performance of nine conical cooling-air ejectors at primary jet pressure ratios from 1 to 10, secondary pressure ratios to 4.0, and a temperature ratio of unity. This phase of the investigation was limited to conical ejectors having shroud exit to primary nozzle exit diameter ratios of 1.06 and 1.40, with several spacing ratios for each. The experimental results indicated that the pumping range and amount of cooling-air flow obtained with a 1.06 diameter ratio ejector were relatively small for cooling purposes but that the maximum possible thrust loss, which occurred with no secondary flow, was only 7 percent of convergent nozzle thrust. The 1.40 diameter ratio ejector produced a large cooling air flow and showed a possible thrust loss of 29.5 percent with no cooling air flow. Thrust gains were attained with ejectors of both diameter ratios at secondary pressure ratios greater than 1.0. The limiting primary pressure ratio below which an ejector can operate at a specific secondary pressure ratio (cut-off point) may be estimated for various flight conditions from data contained herein.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E52F26
    Format: application/pdf
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  • 34
    Publication Date: 2019-06-28
    Description: As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TR-999
    Format: application/pdf
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  • 35
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    In:  CASI
    Publication Date: 2019-06-28
    Description: An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TR-974 , NACA-ACR-L6D02 , NACA-AR-36
    Format: application/pdf
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  • 36
    Publication Date: 2019-08-16
    Description: Contents: Preliminary notes on the efficiency of propulsion systems; Part I: Propulsion systems with direct axial reaction rockets and rockets with thrust augmentation; Part II: Helicoidal reaction propulsion systems; Appendix I: Steady flow of viscous gases; Appendix II: On the theory of viscous fluids in nozzles; and Appendix III: On the thrusts augmenters, and particularly of gas augmenters
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TM-1259
    Format: application/pdf
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  • 37
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    Unknown
    In:  CASI
    Publication Date: 2019-08-13
    Description: The performance of a jet power plant consisting of a compressor and a turbine is determined by the characteristic curves of these component parts and is controllable by the characteristics of the compressor and the turbine i n relation t o each other. The normal. output, overload, and throttled load of the Jet power plant are obtained on the basis of assumed straight-line characteristics.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TM-1258
    Format: application/pdf
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