ALBERT

Description: Propulsion, while conventionally included on the list of important aeronautical disciplines along with aerodynamics, structures, etc., is in itself a systems endeavor, analogous to the engineering of the entire vehicle; indeed propulsion encompasses important aspects of all the other disciplines. In recognition of this fact, the panel focused its discussion on those aspects of the key disciplines that are especially or uniquely important to propulsion. From the initial development of the airplane, the propulsion system has been recognized as one of the pacing technologies. It is perhaps because of the technological disparity between the reciprocating engine and the primitive airframe that the two remained relatively and separate, were developed somewhat independently, usually by different organizations. In recent years, the maturing of the gas turbine power plant and the advance in high-speed airframes have rendered this separation somewhat artificial. The power plant and the airframe now share common structural and aerodynamic elements; as the flight Mach number rises, the degree of interaction increases. By the year 2000, this interdependence will have increased in many respects to a point where independent design may not be practical or possible. During the period since the initiation of the aircraft gas turbine, the solid propellant rocket and the liquid propellant rocket, a vast array of other novel engines have been studied, covering the full spectrum of flight conditions from low subsonic to hypersonic and transatmospheric flight. In each instance, performance limits have been investigated under the assumption that current technology or reasonably foreseeable technology would be available for their development. Among the extensive list of advanced, high-performance concepts and cycles examined are the hypersonic ramjet, the variable cycle, runway-to-orbit airbreathing engine, the ram rocket (airbreathing and rich solid propellant rocket), and the air turborocket. At various times, these systems have come relatively close to meriting development and application. In many instances, limitations of materials and technologies curtailed development. As important and with almost equal frequency, the lack of commercial or military utility of the concept precluded the necessary funding. It is instructive to note that two former items on this list, the turbofan (bypass engine) and the high-speed turboprop, are respectively a mainstay engine and a promising development. In the case of the turbofan, its full potential could not be realized until turbine cooling technology had been developed and new materials developed to permit the construction of transonic fans. In the case of the highspeed turbopropeller engine, not only were the material and turbine technologies needed, but, in addition, the rise in fuel costs provided the impetus to take advantage of its favorable fuel consumption characteristic. As the basic technologies progress and as new missions become attractive, the engines in the foregoing list become candidates for new feasibility studies and further technology development. At the present time, the ram rocket is the prime contender to augment the range of small missiles. Of interest also is the hypersonic ram jet and its logical extension, the runway-to-orbit airbreathing engine. Much of this report deals with the development of current or near-future power plant concepts. First, the motivating factors for aeronautical propulsion research are reviewed as a reminder of the importance of continued effort in a field that has often been characterized as mature. Next, technical areas are discussed in which the panel feels additional research effort is warranted and would lead to the realization of the technological potentials between now and the year 2000. Under these guidelines, new cycles (e.g., isothermal energy exchange) were not considered by the panel. Finally, although facility requirements were not a prime consideration in the current projections, the panel believes that the increasing complexity of propulsion systems; the need for more refined interaction between propulsion system, airframe, and controls; and increasing operation in adverse weather will require test capabilities beyond those now available (see appendix). Enhanced test capability is needed in the areas of propulsion airframe integration and in largescale icing research with proper concurrent treatment of altitude, temperature, and speed.

Description: Results of flow visualization experiments and measurements of the temperature field produced by a single jet and a row of dilution jets issued into a reverse flow combustor are presented. The flow in such combustors is typified by transverse and longitudinal acceleration during the passage through its bending section. The flow visualization experiments were designed to examine the separate effects of longitudinal and transverse acceleration on the jet trajectory and spreading rate. A model describing a dense single jet in a lighter accelerating cross flow is developed. The model is based on integral conservation equations, including the pressure terms appropriate to accelerating flows. It uses a modified entrainment correlation obtained from previous experiments of a jet in a cross stream. The flow visualization results are compared with the model calculations in terms of trajectories and spreading rates. Each experiment is typified by a set of three parameters: momentum ratio, density ratio, and the densimetric Froude number. When injection velocities are large or densities are small, the Froude number becomes very large and hence, unimportant. Therefore, the Froude number is generally significant in describing liquid experiments but is unimportant for the gas experiments in the combustor. Agreement between test and calculated results is encouraging but tends to become poorer with increasing momentum ratio. The temperature measurements are presented primarily in the form of consecutive normalized temperature profiles. Some interpolated isothermal contours are also shown. The single jet trajectories are consistently found to be swept towards the inner wall of the bend, whether injection is from the outer or the inner wall. This behavior is explained by a drifting effect which consists of a transverse velocity component across the combustor due to the developing nature of the flow along it. Plots of lateral temperature distributions of the jet indicate that under longitudinal acceleration conditions the thermal spreading of the jet may be completely suppressed. Comparison between combustor experimental results and model calculations shows poor agreement due to the drifting effect which is not taken into consideration in the model calculations. The row of jets experiments are characterized by two additional parameters: spacing ratio and confinement parameter. The results are shown in the form of consecutive normalized temperature profiles. The confinement parameter appears to become increasingly important with decreasing spacing ratio, in particular when its effect is enhanced by the drifting phenomenon and associated pressure field. A tightly spaced row of jets injected from the inner wall, prior to the bend, is surprisingly kept attached to the inner wall in spite of the strong turning. A similar attachment for a jet injected from the outer wall is not observed.

Description: AiResearch Manufacturing Company, a division of the Garrett Corporation, is engaged in manufacturing a broad variety of products of the aerospace, energy, metals, transit, and marine industries. Products such as fuel controls and rotating accessories for gas turbines are produced by another Garrett division. Computerized analysis is employed to determine how well proposed seals will stand up to stress. For such analysis, AiResearch uses a computer program supplied by COSMIC known as VISCEL. Program saves engineers time in development of software and contributes to improved efficiency in seal analysis.

Description: A semispan wing and nacelle of a typical general aviation twin-engine aircraft was tested to evaluate the cooling capability and drag or several nacelle shapes; the nacelle shapes included cooling air inlet and exit variations. The tests were conducted in the Ames Research Center 40 x 80-ft Wind Tunnel. It was found that the cooling air inlet geometry of opposed piston engine installations has a major effect on inlet pressure recovery, but only a minor effect on drag. Exit location showed large effect on drag, especially for those locations on the sides of the nacelle where the suction characteristics were based on interaction with the wing surface pressures.

Description: Williams International's F107 fanjet engine is used in two types of cruise missiles, Navy-sponsored Tomahawk and the Air Force AGM-86B Air Launched Cruise Missile (ALCM). Engine produces about 600 pounds thrust, is one foot in diameter and weighs only 141 pounds. Design was aided by use of a COSMIC program in calculating airflows in engine's internal ducting, resulting in a more efficient engine with increased thrust and reduced fuel consumption.

Description: A variable geometry gas turbine has an array of ceramic composition vanes positioned by an actuating ring coupled through a plurality of circumferentially spaced turbine vane levers to the outer end of a metallic vane drive shaft at each of the ceramic vanes. Each of the ceramic vanes has an end slot of bow tie configuration including flared end segments and a center slot therebetween. Each of the vane drive shafts has a cross head with ends thereof spaced with respect to the sides of the end slot to define clearance for free expansion of the cross head with respect to the vane and the cross head being configured to uniformly distribute drive loads across bearing surfaces of the vane slot.

Description: A gas turbine engine combustor assembly of annular configuration has outer and inner walls made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween and each outer and inner wall including a transition panel of porous metal defining a combustor assembly outlet supported by a combustor mount assembly including a stiffener ring having a side undercut thereon fit over a transition panel end face; and wherein an annular weld joins the ring to the end face to transmit exhaust heat from the end face to the stiffener ring for dissipation from the combustor; a combustor pilot member is located in axially spaced, surrounding relationship to the end face and connector means support the stiffener ring in free floating relationship with the pilot member to compensate for both radial and axial thermal expansion of the transition panel; and said connector means includes a radial gap for maintaining a controlled flow of coolant from outside of the transition panel into cooling relationship with the stiffener ring and said weld to further cool the end face against excessive heat build-up therein during flow of hot gas exhaust through said outlet.

Description: A gas turbine engine combustor assembly of unique configuration has an outer wall made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween by a reinforcing and heat dissipation ring and a unique weld configuration to prevent thermal erosion of the ends of the porous metal panels at the butt joints; the combustor further including a unique inner wall made up of a plurality of like axially extending multi-layered porous metal panels joined at butt joints by a reinforcing and heat dissipation ring on the inner surface of the inner wall panels and an improved butt weld joint that prevents thermal erosion of the ends of the porous metal inner wall panels.

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.

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,

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.

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.

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.

Description: The performance of a two-stage turbine with variable-area first-stage turbine nozzles was determined in the NACA Lewis altitude wind tunnel over a range of simulated altitudes from 15,000 to 44,000 feet and engine speeds from 50 to 100 percent of rated speed. The variable-area turbine nozzles used in this investigation were primarily a test device for compressor research purposes and were not necessarily of optimum aerodynamic design. The results of this investigation are indicative of effects of turbine-nozzle-area variation on turbine performance within the operating range allowed by the engine. The variable-area turbine nozzles were found to be mechanically reliable and to have negligible leakage losses. Increasing the turbine-nozzle-throat area from 1.15 to 1.67 square feet increased the corrected turbine gas flow or effective turbine nozzle area about 10 percent. At a given corrected turbine speed and turbine pressure ratio, changing the turbine nozzle area from 1.30 to 1. 67 square feet lowered the turbine efficiency 3 or 4 percent. The effect of increasing the turbine nozzle area from 1.15 to 1.67 square feet (decreasing the turning angle about 7 1/2 degrees) would be to lower the turbine efficiency about 5 or 6 percent.

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,

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.

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,

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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 .

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.

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.

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.

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.

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.

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

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.

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.

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.

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.

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.

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.

Description: An altitude-test-chamber investigation was conducted to determine the operational and performance characteristics of a McDonnell afterburner with a fixed-area exhaust nozzle on a J34 engine. At rated engine speed, the altitude limit, as determined by combustion blow-out, occurred as a band of unstable operation of about 6000-foot altitude in width with minimum altitude limits from 31,000 feet at a simulated flight Mach number of 0.40 to about 45,500 feet at a simulated flight Mach number of 1.00. Considerable difficulty was experienced in attempting to establish or maintain balanced-cycle engine operation at altitudes above 36,000 feet. The fuel-air ratio for balanced-cycle operation and lean blowout of the afterburner, the augmented-thrust ratio, the total specific fuel consumption, and the afterburner combustion efficiency for balanced-cycle operation are summarized in a table. Satisfactory afterburner ignition was obtained over a range of flight Mach Numbers from 0.32 to 0.60 at altitudes from 10,000 to 30,000 and engine speeds from 10,000 to 12,500 rpm.

Description: A method for calculation of a counterrotating propeller which is similar to Walchner's method for calculation of the single propeller in the free air stream is developed and compared with measurements. Several dimensions which are important for the design are given end simple formulas for the gain in efficiency derived. Finally a survey of the behavior of the propeller for various operating conditions is presented.

Description: Results of measurements on a shrouded propeller are given. The propeller is designed for the high ratio of advance and high thrust loading. The effect of the shape of propeller and shroud upon the aerodynamic coefficients of the propulsion unit can be seen from the results. The highest efficiency measured is 0.71. The measurements permit the conclusion that the maximum efficiency can be essentially improved by shroud profiles of small chord and thickness. The largest static thrust factor of merit measured reaches according to Bendemann, a value of about zeta = 1.1. By the use of a nose split flap the static thrust for thin shroud profiles with small nose radius can be about doubled. In a separate section numerical investigations of the behavior of shrouded propellers for the ideal case and for the case with energy losses are carried out. The calculations are based on the assumption that the slipstream cross section depends solely on the shape of the shroud and not on the propeller loading. The reliability of this hypothesis is confirmed experimentally and by flow photographs for a shroud with small circulation. Calculation and test are also in good agreement concerning efficiency and static thrust factor of merit. The prospects of applicability for shrouded propellers and their essential advantages are discussed.

Description: The requirements on gas turbines for aircraft power units, namely, adequate efficiency, operation at high gas temperatures, low weight, and small dimensions, must be taken into consideration during the design of the blading. To secure good efficiency, it is necessary that the gas flow past the blades as smoothly as possible without separation. This is relatively easily obtainable in the accelerated flow of turbine blading, if the blade spacing is chosen small enough. A small blade spacing, however, is detrimental to the other requirements outlined above. Operation at high gas temperatures usually calls for blade cooling. This cooling is associated with a power input that lowers the turbine efficiency. Since the amount of heat that must be carried off for coding a blade can be influenced rather little, the gross power input for a turbine stage can be reduced by keeping the number of blades to a minimum, that is, with blades of high spacing ratio. But here also a limit is imposed, the exceeding of which is followed by separation of flow. Hence the requirement of finding blade forms on which the flow separates at rather high spacing ratios .

Description: An investigation was conducted to determine the performance characteristics of the rotor and inlet guide vanes used in the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. Outlet stators used in the engine were omitted to facilitate study of the supersonic rotor. The extent of the deviation from design performance indicates that the design-shock configuration was not obtained. A maximum pressure ratio of 2.26 was obtained at an equivalent tip speed of 1614 feet per second and an adiabatic efficiency of 0.61. The maximum efficiency obtained was 0.79 at an equivalent tip speed of 801 feet per second and a pressure ratio of 1.29. The performance obtained was considerably below design performance. The effective aerodynamic forces encountered appeared to be large enough to cause considerable damage to the thin aluminum leading edges of the rotor blades.

Description: As part of the performance investigation of compressors for the J33 turbojet engine, the A-21 model and the A-23 model with a 17- and a 34-blade impeller were operated with water injection at their respective design equivalent speeds of 11,500 and 11,750 rpm. Inlet conditions of pressure of 14 inches of mercury absolute and of ambient temperature correspond to those of the investigation of these models without water injection. The water-air ratio by weight ranged from 0.05 to 0.06. By the use of water injection, the peak pressure ratio of the A-21 compressor and the A-23 compressor with a 34-blade impeller increased approximately 0.38, whereas that of the A-23 compressor with a 17-blade impeller increased only 0.14. The decrease in maximum efficiency for the three compressors ranged from 0.12 to 0.14. The highest increase in maximum equivalent weight flow of air plus weight flow of water was 10.90 pounds per second obtained with the A-21 compressor. The increase in air weight flow alone was approximately 5.70 pounds per second for the A-21 compressor end the A-23, 17-blade compressor, which exceeded the increase of 3.15 pounds per second for the A-23; 34-blade compressor.

Description: A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzle designs in combination with special rotor blades having a 20 inlet angle. The performance is presented in terms of blade, rotor, and brake efficiency as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. The blade efficiency with the nozzle having circular pas- sages (K) was equal to or higher than that with the nozzle having rectangular passages (J) for all pressure ratios and speeds investigated. The maximum blade efficiency of 0.571 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.296. The difference in blade efficiency was negligible at a pressure ratio of 8 at the low speeds; the maxim difference was 0.040 at a pressure ratio of 20 and a blade-jet speed ratio of 0.260.

Description: The J33-A-27 compressor was operated at an inlet pressure of 14 inches of mercury absolute and ambient inlet temperature over a range of equivalent impeller speeds from 6100 to 11,800 rpm. At the design equivalent speed of 11,800 rpm, the J33-A-27 compressor had a peak pressure ratio of 4.40 at an equivalent weight flow of 105.7 pounds per second and a peak adiabatic temperature-rise efficiency of 0.745. The maximum equivalent weight flow at design speed was 113.5 pounds per second.

Description: An investigation was conducted to determine the effect of turbine-disk cooling with air on the efficiency and the power output of the radial-flow turbine from the Turbo Engineering Corporation TT13-18 turbosupercharger. The turbine was operated at a constant range of ratios of turbine-inlet total pressure to turbine-outlet static pressure of 1,5 and 2.0, turbine-inlet total pressure of 30 inches mercury absolute, turbine-inlet total temperature of 12000 to 20000 R, and rotor speeds of 6000 to 22,000 rpm, Over the normal operating range of the turbine, varying the corrected cooling-air weight flow from approximately 0,30 to 0.75 pound per second produced no measurable effect on the corrected turbine shaft horsepower or the turbine shaft adiabatic efficiency. Varying the turbine-inlet total temperature from 12000 to 20000 R caused no measurable change in the corrected cooling-air weight flow. Calculations indicated that the cooling-air pumping power in the disk passages was small and was within the limits of the accuracy of the power measurements. For high turbine power output, the power loss to the compressor for compressing the cooling air was approximately 3 percent of the total turbine shaft horsepower.

Description: An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ-55-FF-1 turbo Jet engine. The test unit consisted of a row of inlet guide vanes and a supersonic rotor; the stator vanes after the rotor were omitted. The maximum pressure ratio produced in the single stage was 2.28 at an equivalent tip speed or 1814 feet per second with an adiabatic efficiency of approximately 0.61, equivalent weight flow of 13.4 pounds per second. The maximum efficiency of 0.79 was obtained at an equivalent tip speed of 801 feet per second.

Description: An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. An analysis of the performance of the rotor was made based on detailed flow measurements behind the rotor. The compressor apparently did not obtain the design normal-shock configuration in this investigation. A large redistribution of mass occurred toward the root of the rotor over the entire speed range; this condition was so acute at design speed that the tip sections were completely inoperative. The passage pressure recovery at maximum pressure ratio at 1614 feet per second varied from a maximum of 0.81 near the root to 0.53 near the tip, which indicated very poor efficiency of the flow process through the rotor. The results, however, indicated that the desired supersonic operation may be obtained by decreasing the effective contraction ratio of the rotor blade passage.

Description: A theory has been developed for resetting the blade angles of an axial-flow compressor in order to improve the performance at speeds and flows other than the design and thus extend the useful operating range of the compressor. The theory is readily applicable to the resetting of both rotor and stator blades or to the resetting of only the stator blades and is based on adjustment of the blade angles to obtain lift coefficients at which the blades will operate efficiently. Calculations were made for resetting the stator blades of the NACA eight-stage axial-flow compressor for 75 percent of design speed and a series of load coefficients ranging from 0.28 to 0.70 with rotor blades left at the design setting. The NACA compressor was investigated with three different blade settings: (1) the design blade setting, (2) the stator blades reset for 75 percent of design speed and a load coefficient of 0.48, and (3) the stator blades reset for 75 percent of design speed and a load coefficient of 0.65.

Description: A wind tunnel investigation was conducted to determine the performance of a 4000-pound-thrust axial-flow turbojet engine with a high flow compressor. Pressure altitudes included 5000 to 40000 feet with ram pressure ratios from 1.00 to 1.82. Altitudes included 20000 to 40000 feet and ram pressure ratios from 1.09 to 1.75. A comparison is made between engine performance with high flow and low flow compressors.

Description: A wind tunnel investigation was conducted to determine the performance of a turbine operating as an integral part of a turbojet engine. Data was obtained while the engine was running over full operable range of speeds at various altitudes and flight mach numbers, and with four nozzles of different outlet areas.A maximum turbine efficiency of 0.875 was obtained at altitude of 15 thousand feet, Mach number 0.53, and corrected turbine speed of 5900 rpm.

Description: Temperature and pressure distributions for an original and modified 3000 pound thrust axial flow turbojet engine were investigated. Data are included for a range of simulated altitudes from 5000 to 45000 feet, Mach numbers from 0.24 to 1.08, and corrected engine speeds from 10,550 to 13,359 rpm.

Description: This report presents the results of the tests of a power-plant installation to improve the circumferential pressure-recovery distribution at the face of the engine. An underslung "C" cowling was tested with two propellers with full cuffs and with a modification to one set of cuffs. Little improvement was obtained because the base sections of the cuffs were stalled. A set of guide vanes boosted the over-all pressures and helped the pressure recoveries for a few of the cylinders. Making the underslung cowling into a symmetrical "C" cowling evened the pressure distribution; however, no increases in front pressures were obtained. The pressures at the top cylinders remained low and the high pressures at the bottom cylinders were reduced. At higher powers and engine speeds, the symmetrical cowling appeared best from the standpoint of over-all cooling characteristics.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of an axial flow-type turbojet engine with a 4000-pound-thrust rating over a range of pressure altitudes from 5,000 to 50,OOO feet, ram pressure ratios from 1.00 to 1.86, and temperatures from 60 deg to -50 deg F. The low-flow (standard) compressor with which the engine was originally equipped was replaced by a high-flow compressor for part of the investigation. The effects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, acceleration, starting, operation of fuel-control systems, and bearing cooling were investigated. With the low-flow compressor, the engine could be operated at full speed without serious burner unbalance at altitudes up to 50,000 feet. Increasing the altitude and airspeed greatly reduced the operable speed range of the engine by raising the minimum operating speed of the engine. In several runs with the high-flow compressor the maximum engine speed was limited to less than 7600 rpm by combustion blow-out, high tail-pipe temperatures, and compressor stall. Acceleration of the engine was relatively slow and the time required for acceleration increased with altitude. At maximum engine speed a sudden reduction in jet-nozzle area resulted in an immediate increase in thrust. The engine started normally and easily below 20,000 feet with each configuration. The use of a high-voltage ignition system made possible starts at a pressure altitude of 40,000 feet; but on these starts the tail-pipe temperatures were very high, a great deal of fuel burned in and behind the tail-pipe, and acceleration was very slow. Operation of the engine was similar with both fuel regulators except that the modified fuel regulator restricted the fuel flow in such a manner that the acceleration above 6000 rpm was very slow. The bearings did not cool properly at high altitudes and high engine speeds with a low-flow compressor, and bearing cooling was even poorer with a high-flow compressor.

Description: The effect of rotor-blade length, inlet angle, and shrouding was investigated with four different nozzles in a single-stage modification of the Mark 25 aerial-torpedo power plant. The results obtained with the five special rotor configurations are compared with those of the standard first-stage rotor with each nozzle. Each nozzle-rotor combination was operated at nominal pressure ratios of 8, 15 (design), and 20 over a range of speeds from 6000 rpm to the design speed of 18,000 rpm. Inlet temperature and pressure conditions of 1OOOo F and 95 pounds per square inch gage, respectively, were maintained constant for all runs.

Description: Flow-metering devices used by the NACA and by the manufacturer of the J33 turbojet engine were calibrated together to determine whether an observed discrepancy in weight flow of approximately 4 percent for the two separate investigations might be due to the different devices used to meter air flow. A commercial adjustable orifice and a square-edge flat-plate orifice used by the NACA and a flow nozzle used by the manufacturer were calibrated against surveys across the throat of the nozzle. It was determined that over a range of weight flows from 18 to 45 pounds per second the average weight flows measured by the metering device used for the compressor test would be 0.70 percent lower than those measured by the metering device used in the engine tests and the probable variation about this mean would be +/- 0.39 percent. The very close agreement of the metering devices shows that the greater part of the discrepancy in weight flow is attributable to the effect of inlet pressure.

Description: An investigation was conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of the X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet,simulated flight Mach numbers from 0 to 1.08, and engine speeds from 4000 to 12,500 rpm. Performance data are presented to show graphically the effects of altitude at a flight Mach number of 0.25 and of flight Mach number at an altitude of 25,000 feet. The performance data are generalized to show the applicability of methods used to determine performance at any altitude from data obtained at a given altitude. A complete tabulation of performance data, as well as lubrication- and fuel- system data, is presented.

Description: Investigations were made of the turbine from a Mark 25 torpedo to determine the performance of the unit with three different turbine nozzles at various axial nozzle-wheel clearances. Turbine efficiency with a reamed nondivergent nozzle that uses the axial clearance space for gas expansion was little affected by increasing the axial running clearance from 0.030 to 0.150 inch. Turbine efficiency with cast nozzles that expanded the gas inside the nozzle passage was found to be sensitive to increased axial nozzle-wheel clearance. A cast nozzle giving a turbine brake efficiency of 0.525 at an axial running clearance of 0.035 inch gave a brake efficiency of 0.475 when the clearance was increased to 0.095 inch for the same inlet-gas conditions and blade-jet speed ratio. If the basis for computing the isentropic power available to the turbine is the temperature inside the nozzle rather then the temperature in the inlet-gas pipe, an increase in turbine efficiency of about 0.01 is indicated.

Description: At the request of the Air Material Command, Arm Air Forces, an investigation was conducted at the NACA Cleveland laboratory to determine the performance characteristics of the XJ-41-V turbojet-engine compressor. The complete compressor was mounted on a collecting chamber having an annular air-flow passage simulating the burner annulus of the engine and was driven by an electric motor. The compressor was extensively instrumented to determine the overall performance of the compressor, the characteristic performance of each of the compressor components, the state of the air stream in the simulated burner annulus, and the operation of the compressor bearings. An initial investigation at an equivalent compressor speed of 8000 rpm was made to determine the performance of the compressor and the collecting chamber and to determine the similarity of the air stream at the entrance to the simulated burner annulus. The mechanical performance of the compressor over a range of actual compressors speeds from 3300 to 8000 rpm is reported.

Description: Measurements on three tubes with flow regulated by suction at the trainling edge of the tube are described. It was possible to vary the mass of air flowing through the tube over a large range. Such tubes could be used for shrouded propellers.

Description: A preliminary investigation of an axial-flow gas turbine-propeller engine was conduxted. Performance data were obtained for engine speeds from 8000 to 13,000 rpm and altitudes from 5000 to 35,000 feet and compressor inlet ram pressure ratios from 1.00 to 1.17.

Description: A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

Description: Operating characteristics of the 11-stage 4000-pound-thrust axial-flow turbojet engine were determined. A standard compressor and a compressor with the blade angles of the rotor and stator blades increased 5 degrees to obtain greater air flow, were investigated.

Description: Combustion chamber performance properties of a 3000-pound-thrust axial-flow turbojet engine were determined. Data are presented for a range of simulated altitudes from 15,000 to 45,0000 feet and a range of Mach numbers from 0.23 to 1.05 for various modifications of the engine.

Description: An investigation was conducted in the NACA Cleveland altitude wind tunnel to determine the operational characteristics of the Westinghouse 19B-2, 19B-8, and 19XB-l jet-propulsion engines. The 19B engine is one af the earliest experimental Westinghouse axial flow engines. The 19XB-1 engine is an experimental prototype of the Westinghouse 15 series, having a rated thrust of 1400 pounds. Improvements in performance and operational characteristics have resulted in the 19XB-2B engine with a rated thrust of 1600 pounds. The operational characteristics were determined over a range of simulated altitudes from 5000 to 30,000 feet for the 19B engines and from 5000 to 35000 feet for the 19XB-l engine at airspeed from 20 to 380 miles per hour. The affects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, starting, acceleration, and functioning of the fuel-control system are discussed. Damage to the engines that occurred during the investigation is also briefly discussed. The changes made in the combustion-chamber configuration to improve the operating we are described.

Description: A theoretical investigation has been made of various methods of thrust augmentation for turbojet engines. The method investigated were tail-pipe burning, water injection at the compressor inlet, a combination of tail-pipe burning and water injection, bleedoff in conjunction with water injection at the compressor inlet, and rocket assist. The effect of ratio of augmented-to-normal total liquid consumption, flight conditions, and design compressor pressure ratio on the augmentation produced by each method were determined. A comparison was also made for a given time of operation of the weight of an augmented engine plus fuel and additional liquids to the weight of a standard engine plus fuel producing the same thrust.

Description: The Allison model 400-C6 compressor was operated at an inlet pressure of 12 inches of mercury absolute ana ambient inlet temperature at equivalent impeller speeds of 6000, 7000, and 8500 rpm. Additional runs at an equivalent speed of 7000 rpm and ambient inlet temperature were made at inlet pressures from 7 to 22 inches of mercury absolute. The results of this investigation are compared with those of the 533-A-23 compressors. For the speeds investigated, the Allison model 400-C6 compressor had a maximum adiabatic temperature-rise efficiency of 0.768 at an equivalent speed of 7000 rpm; the corresponding equivalent weight flow was 45.0 pounds per second and the pressure ratio was 1.83. At an equivalent impeller speed of 8500 rpm, the maximum equivalent weight flow was 61.6 pounds per second and the peak pressure ratio of 2.38 occurred at an equivalent weight flow of 52.2 pounds per 1 second and an adiabatic temperature-rise efficiency of 0.714. At an equivalent speed of 7000 rpm, increasing the compressor- inlet pressure increased the maximum equivalent weight flow and the pressure ratio.

Description: The production-model 333-A-23 turbojet-engine compressor with a 17-blade impeller was operated at ambient and 0 F inlet temperatures and at inlet pressures of 14 and 5 inches mercury absolute for equivalent impeller speeds from 6000 to 12,750 rpm. The results of this investigation are compared with those of the 533-A-21 compressor. At the design equivalent speed of 11,750 rpm the maximum pressure ratio was 4.39. This occurred at the surge point at which the equivalent weight flow was 80.8 pounds per second, ana the adiabatic temperature-rise efficiency was 0.757. The maximum flow at the design equivalent speed was 88.0 pounds per second. The maximum adiabatic temperature-rise efficiency of 0.799 was obtained at an equivalent speed of 10,000 rpm, and equivalent weight flow of 62.9 pounds per second, and a pressure ratio of 3.20. At the maximum equivalent speed investigated (12,750 rpm), a peak pressure ratio of 4.90 was attained at an equivalent weight flow of 85.4 pounds per second and an efficiency of 0.680.

Description: In an investigation of the J-33-A-21 and the J-33-A-23 compressors with and without water injection, it was discovered that the compressors reacted differently to water injection although they were physically similar. An analysis of the effect of water injection on compressor performance and the consequent effect on matching of the compressor and turbine components in the turbojet engine was made. The analysis of component matching is based on a turbine flow function defined as the product of the equivalent weight flow and the reciprocal of the compressor pressure ratio.

Description: An investigation has been conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance and windmilling drag characteristics of an original and a modified turbojet engine of the same type. Data have been obtained at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.09 to 1.08, and engine speeds from 4000 to 12,500 rpm. Engine performance data are presented for both engines to show the effects of altitude at a flight Mach number of 0.25 and of flight Mach number at an altitude of 25,000 feet. Performance of the original and modified engines is compared for a range of simulated flight conditions. The performance data are generalized to show the applicability of methods used to estimate performance at any altitude from data obtained at a given altitude. Engine-windmilling-speed and windmilling-drag data are presented for a range of simulated flight conditions.

Description: An investigation was conducted in an altitude test chamber to determine the effects of inlet airflow distortion on the compressor steady-state and surge characteristics of a high-pressure ratio, axial-flow turbojet engine. Circumferential-type inlet flow distortions were investigated, which covered a range of distortion sector angles from 20 deg to 168 deg and distortion levels up to 22 percent. The presence of inlet airflow distortions at the compressor face resulted in a substantial increase in the local pressure ratio in the distorted region, primarily for the inlet stages. The local pressure ratio in the distorted region for the inlet stages increased as either the distortion sector angle decreased or the percent distortion increased. The average compressor-surge pressure ratio was much more sensitive to inlet airflow distortions at lower engine speeds than at engine speeds near rated. Hence, compressor-surge margin reduction due to inlet airflow distortion was quite severe at the lower engine speeds. Although the average compressor-surge pressure ratio was generally reduced with inlet flow distortion, local pressure ratios across the distorted sector of the compressor were obtained during surge and were significantly greater than the normal compressor-surge pressure ratio. This was a result of increased loading of the inlet stages in the distorted region.

Description: An altitude-test-chamber investigation was conducted to determine the operational characteristics and altitude blow-out limits of a Solar afterburner in a 24C engine. At rated engine speed and maximum permissible turbine-discharge temperature, the altitude limit as determined by combustion blow-out occurred as a band of unstable operation of about 8000 feet altitude in width with maximum altitude limits from 32,000 feet at a Mach number of 0.3 to about 42,000 feet at a Mach number of 1.0. The maximum fuel-air ratio of the afterburner, as limited by maximum permissible turbine-discharge gas temperatures at rated engine speed, varied between 0.0295 and 0.0380 over a range of flight Mach numbers from 0.25 to 1.0 and at altitudes of 20,000 and 30,000 feet. Over this range of operating conditions, the fuel-air ratio at which lean blow-out occurred was from 10 to 19 percent below these maximum fuel-air ratios. Combustion was very smooth and uniform during operation; however, ignition of the burner was very difficult throughout the investigation. A failure of the flame holder after 12 hours and 15 minutes of afterburner operation resulted in termination of the investigation.

Description: With the further development of axial blowers into highly loaded flow machines, the influence of the diameter ratio upon air output and efficiency gains in significance. Clarification of this matter is important for single-stage axial compressors, and is of still greater importance for multistage ones, and particularly for aircraft power plants. Tests with a single-stage axial blower gave a decrease in the attainable maximum pressure coefficient and optimum efficiency as the diameter ratio increased. The decrease must be ascribed chiefly to the guide surface of the hub and housing between the blades increasing with the diameter ratio.

Description: As part of an investigation af the application of nuclear energy to various types of power plants for aircraft, calculations have been made to determine the effect of several operating conditions on the performance of condensers for mercury-turbine power plants. The analysis covered 8 range of turbine-outlet pressures from 1 to 200 pounds per square inch absolute, turbine-inlet pressures from 300 to 700 pounds per square inch absolute,and a range of condenser cooling-air pressure drops, airplane flight speeds, and altitudes. The maximum load-carrying capacity (available for the nuclear reactor, working fluid, and cargo) of a mercury-turbine powered aircraft would be about half the gross weight of the airplane at a flight speed of 509 miles per hour and an altitude of 30,000 feet. This maximum is obtained with specific condenser frontal areas of 0.0063 square foot per net thrust horsepower with the condenser in a nacelle and 0.0060 square foot per net thrust horsepower with the condenser submerged in the wings (no external condenser drag) for a turbine-inlet pressure of 500 pounds per square inch absolute, a turbine-outlet pressure of 10 pounds per square inch absolute, and 8 turbine-inlet temperature of 1600 F.

Description: The J33-A-23 compressor with a 34-blade impeller was operated at ambient inlet temperature and an inlet pressure of 14 inches mercury absolute over a range of equivalent impeller speeds from 6000 to 11,750 rpm. Additional runs at equivalent speeds of 7,000, 10,000, and 11,750 rpm and ambient inlet temperature were made at inlet pressures of 5 and 10 inches mercury absolute. The results of this investigation are compared with those of the J33-A-23 compressor with a 17-blade impeller. At the design equivalent speed of 11,750 rpm the 533-A-23 compressor with a 34-blade impeller had a peak pressure ratio of 4.49 at an equivalent weight flow of 82.4 pounds per second and an adiabatic temperature-rise efficiency of 0.740. The maximum equivalent flow at design speed was 91.8 pounds per second. The peak efficiency at design speed (0.757) occurred at an equivalent weight flow of 85.5 pounds per second. The maximum adiabatic temperature- rise efficiency of 0.773 was obtained at an equivalent impeller speed of 10,000 rpm, an equivalent weight flow of 65.8 pounds per second, and a pressure ratio of 3.27. At equivalent impeller speeds of.l0,000 and 11,75O rpm a decrease in inlet pressure resulted in a decrease in maximum equivalent weight flow, peak pressure ratio, and peak adiabatic temperature- rise efficiency.

Description: An investigation of the XJ-41-V turbojet-engine compressor was conducted to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal-type compressors. The results of the research conducted on the original compressor indicated the compressor would not meet the desired engine-design air-flow requirements because of an air-flow restriction in the vaned collector. The compressor air-flow choking point occurred near the entrance to the vaned-collector passage and was instigated by a poor mass-flow distribution at the vane entrance and from relatively large negative angles of attack of the air stream along the entrance edges of the vanes at the outer passage wall and large positive angles of attack at the inner passage wall. As a result of the analysis, a design change of the vaned collector entrance is recommended for improving the maximum flow capacity of the compressor.

Description: The performance of an annular combustion chamber from a 24C turbojet engine was investigated over a range of simulated altitudes from 20,000 to 55,000 feet and corrected engine rotor speeds from 6000 to 13,000 rpm at a simulated ram-pressure ratio of 1.04. The purpose of the investigation was to determine the effects on the altitude operational limits, combustor-outlet gas temperature distribution, combustion efficiencies, and combustor inlet-to-outlet total-pressure drops of two changes in the 24C-4B basket air-passage arrangements that were designed to improve combustor-outlet temperature distribution. These changes were: (a) replacement of the downstream secondary air holes with large rectangular slots further upstream (rectangular-slot basket), and (b) enlargement of anticoking holes in the rectangular-slot basket (modified rectangular-slot basket). The results indicate that improved outlet-gas temperature distribution of each succeeding combustor basket investigated was attained at a sacrifice in the altitude limit of operation. The altitude limits of operation of the combustor with the original basket ranged from 34,000 feet at a corrected engine speed of 6000 rpm to a maximum of 52,000 feet at 12 ' 500 rpm. The altitude limits of the rectangular-slot basket were about 2000 feet lower throughout the engine speed range than those of the original basket. The altitude limits of the combustor with the modified rectangular-slot basket were about equivalent to those of the other baskets in the corrected-engine-speed range from 12,000 to 12,500 rpm but were about 10,000 feet lower than those of the original basket in the corrected-engine-speed range from 6000 to 9000 rpm. For the same inlet-air conditions, the combustion efficiencies were highest for the original basket and progressively lower for each of the other two baskets. The combustor inlet-to-outlet pressure drops of all three combustor baskets at the same operating conditions were within +/- 10 percent of the pressure drop of the original basket.

Description: Compressor operation at low air flows for a given speed is limited by unstable flow conditions, commonly called surge. An investigation of surge in centrifugal compressors (reference 1) showed that the pulsation of pressures and velocities occurred when the slope of the compressor characteristic curve was positive and that the magnitude and frequency, as well as the incidence of surge, depended on the capacity and resistance of the total system. Although the theory presented in reference 1 is applicable to axial-floe compressors, little experimental information is available on the surge characteristics of the individual stages of axial-flow compressors, or on the variation of the surge characteristics with operating conditions. During the investigation to determine the performance of the X24C-2 compressor (references 2 and 3), instrumentation was added to study the surge characteristics and to determine the effect of speed and inlet pressure on the frequency, amplitude, and phase relation of the pressure pulsations behind each stage.

Description: Compressor performance properties for two 11-stage compressors of 3000-pound-thrust axial-flow turbojet engines were determined. Data are presented for a range of simulated altitudes and a range of Mach numbers for various modifications of the engine.

Description: Wind tunnel investigations were performed to determine the performance properties of an axial-flow gas turbine-propeller engine II. Windmilling characteristics were determined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Description: A simulated altitude performance of a 25 1/2-inch-diameter annular-type turbojet combustor was performed to determine the effect of the distribution of basket-hole area on the altitude operational limits of the engine as imposed by the combustor.Total pressure drop was recorded, as well as the effect of fuel-nozzle flow capacity,and fuel-nozzle spray angle for one basket configuration. General observations were made for all configurations regarding flames, extent of afterburning, and durability of the baskets.

Description: An investigation was conducted to evaluate the operational characteristics of a 3000 pound thrust axial flow turbojet engine over a range of simulated altitudes from 2000 to 50,000 feet and simulated flight Mach numbers from 0 to 1.04 throughout the operable range of engine speeds. Engine operating range, acceleration, deceleration, starting, altitude, and flight Mach number compensation of the fuel control system, and operation of the lubrication system at high and low ambient air temperatures were evaluated.

Description: An investigation is being conducted to determine the altitude performance characteristics of the Nene II engine and its components. The present paper presents preliminary results obtained using a jet nozzle of 18.41 inches in diameter, giving an area equal to 96.4 percent of the area of the standard jet nozzle of this engine. The test results presented are for conditions simulating altitudes from seal level to 50,000 feet and ram-pressure ratios from 1.00 to 2.70. The ram pressure ratios correspond to flight Mach numbers between zero and 1.28.

Description: At the request of the Air Material Command, Army Air Forces, an investigation was conducted by the NACA Cleveland laboratory to determine the performance characteristics of the compressor of the XJ-41-V turbojet engine. This report is the second in a series presenting the compressor performance and analysis of flow conditions in the compressor. The static-pressure variation in the direction of flow through the compressor and the location and the cause of the maximum flow restriction at an equivalent speed of 8000 rpm are presented. After the initial runs were reported, the leading edges of the impeller blades and the diffuser surfaces were found to have been roughened by steel particles from a minor failure of auxiliary equipment. The leading edges of the impeller blades were refinished and all high spots resulting from scratches in the diffuser and the accessible parts of the vaned collector passages were removed. The initial overall performance and that obtained with the refinished blades are presented.

Description: An extended analysis was made of the previously reported performance investigation of the original compressor from the XJ-41-v turbojet engine and a similar compressor revised a to obtain a 33-percent increase in the geometric passage area at the vaned-collector entrance. This analysis was based on the concept of the vaned-collector entrance as the throat section of a nozzle. Because of nonuniform air distribution at the vaned-collector entrance, approximately 90 percent of the available flow area was utilized in the original compressor and 94percent in the revised com$ressor. The increase in maximum weight flow obtained with the revised compressor was disproportionate to the increased effective critical throat area because. the air density at the revised vaned-collector entrance for maximum flow was lower than that obtained in the original compressor. This reduction in density resulted from the large pressure losses near the impeller inlet of the revised compressor, which is indicative of impending flow choking in the impeller, The.calculated maximum corrected weight-flow capacity of a compressor consisting of the revised vaneless diffuser and vaned collector with a theoretical impeller that combined peak impeller pressure ratio and peak impeller efficiency at the . maximum flow point would be 112 pounds per second for an equivalent impeller speed of 11,500 rpm;

Description: Performance properties and operational characteristics of an axial-flow gas turbine-propeller engine were determined. Data are presented for a range of simulated altitudes from 5,000 to 35,0000 feet, compressor inlet- ram pressure ratios from 1.00 to 1.17, and engine speeds from 8000 to 13,000 rpm.

Description: Charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet engine for any given set of operating conditions and component efficiencies. The effects of the pressure losses in the inlet duct and combustion chamber, the variation in the physical properties of the gas as it passes through the cycle, and the change in mass flow by the addition of fuel are included. The principle performance charts show the effects of the primary variables and correction charts provide the effects of the secondary variables.

Description: The performance of hypothetical turbojet systems, without thrust augmentation, as power plants for supersonic airplanes has been calculated. The thrust, thrust power, air-fuel ratio, 1 specific fuel consumption, cross-sectional area, and thrust coefficient are shown for free-stream Mach numbers from 1.2 to 3. For comparison, the performance of ram-jet systems over the same Mach number range has also been calculated. For Mach numbers between 1.2 and 2 the calculated thrust coefficient of the turbojet system was found to be larger than the estimated drag coefficient, and the specific fuel consumption was calculated to be considerably less than the specific fuel consumption of the ram-jet system. The turbojet system therefore appears to merit consideration as a propulsion method for free-stream Mach numbers between approximately 1.2 and 2.

Description: Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

Description: A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: A theoretical analysis of the cross-sectional temperature distribution of a water-cooled turbine blade was made using the relaxation method to solve the differential equation derived from the analysis. The analysis was applied to specific turbine blade and the studies icluded investigations of the accuracy of simple methods to determine the temperature distribution along the mean line of the rear part of the blade, of the possible effect of varying the perimetric distribution of the hot gas-to -metal heat transfer coefficient, and of the effect of changing the thermal conductivity of the blade metal for a constant cross sectional area blade with two quarter inch diameter coolant passages.